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^ U S U V L A B S T E C H N I C A L R E P O RT 6 7 * 5

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I I . S . I I H I Y » Y U T I O N M AT E R I E L L A B O R AT O R I E S

F O RT E I I S T I S , V I R G i l A

CONTRACT DA-44-177-AMC-154(T)

PiASECKI AIRCRAFT CORPORATION

PHILADELPHIA, PENNSYLVANIA

L Thi* iocumfni as rrn approvfd

f^r public elease otui ale; us

ttri/)ü« n s nlimited.

JV

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■

DEPARTMENT OF THE ARMY U >. AMMY AVIATIOH MATCfMCl . LABCIUTOMCS

FOOT C U S T S . VIRCINI* 2>*04

The 16H-1A f l ight tes t research effort was conducted as part of th e exploratory development program in support of AAFSS (Advanced Aerial Fire Suppor t System). hi s effort wa s Ini t i - ated in May 1964, an d f l ights of th e 16H-IA aircraft were ended in July 1966. he object ive of th e program wa s to obtain f l ight tes t data at speeds up to 200 knots for future hlgh-perfonaance rotary-wing aircraft designs . he 16H-1A at tained a maxiinum true airspeed in level f l ight of 16 0 knots a t sea-level standard-day cond i t ions and at a gross weight of approximately 6200 pounds. This speed performance wa s less than ant icipated; consequent ly, use of th e program resul ts for high-speed des ign cri ter ia is limited.

Tais report is publ i shed to release th e data obtained under Ch« contract. urther analysis or interpretat ion of th e resul t« contained in this report is subject to th e l imi t a t ions an d scat ter of the data obta ined under the program. he analyses , conclusions , an d reconaendatlons cont i ined herein ar e those of . the invest igator and ar e not confirmed or endorsed by the Government through the publ icat ion of this report .

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*

Di»cUi»er»

When Governaent drawings, specificat ions, or other data ar e used for any purpose other than in connection with a def in i te ly related Govenwent procuroKcnt operation, the United States Government thereby incurs no responsibil i ty nor any obligation whatsoever; an d the fact that the Govern-

•Mt My have fonu la ted , furnished, or i n any way supplied the said drawings, specifications, or other data is not to be regarded by impli- cation or otherwise as in any Banner l icensing the holder or any other person or corporat ion, or conveying an y r ights or pcraission, to manufacture, use, o r sel l any patented invention that may in any wa y be related thereto.

Trade names cited in this report do not consti tute an official emiorse- B t e n t or approval of the use of such coaaercial hardware or software.

vm

Disposit ion Instructions

Destroy this report when no longer needed. o not return i t t o originator.

nm KCTW H

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·•·

THIS DOCUMENT IS BEST

QUALITY AVAILABLE TH COPYFURNISHED TO DTIC CONTAINED

SIGNIFICANT NUMBER OF

PAGES WHICH DO NOTREPRODUCE LEGI LYo

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*> Task 1F162203A14311

Contract A-iJil-lTT-AMC-lS^CT)

ÜSAAVLAüS echnical eport 7-58

August 968

16K-1A FLIGHT TEST RESEARCH PROGRAM

Final Repor t

by

D.N. Meyers , L . V. Tompklns , J .H. G oldbe rg

Prepa red by

Plaseck l Aircraf t C o rpo ra t ion

Philadelphia , Pennsy lvan ia

f o r

U. S . ARMY AVIATION MATERIEL LABORATORIES

PORT EUSTIS, VIRGINIA

This document has been approved

for public elease and ale; ts

distribution s unlimited.

» ii miiiw—MiiiiiwiMiiiiriMwiMW n w r

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SUMMARY

This eport resents he esults f light est rogram conducted y iaseckl ircraft orporation f he iaseckl M odel 6H-1A haft-driven ompound elicopter. The lying qualities nd erforaance f he 6H-1A ircraft ere nvesti-

gated ver evel-flight peed ange f o 67 nots nd dive peeds p o 95 nots. These ests esulted n etter understanding f he ndependent esign arameters n he om - pound helicopter, uch s he all ropeiier/main otor power distribution nd he istribution f ift etween he ing nd main otor.

The nterrelationship

f

he xea

f

ift,

hrust, rag,

and ngular ttitudes f he ing, fuselage, nd otor xis had een nvestigatei reviously uring ore han 0 0 ours f ground nd light esting f M otael 6H-i t The odel PH^IA

test rogram has rovided ecorded easurements f hese parameters« In btainin.5 hese ata, he 6ii-1Ä ade 71 lights and ccumulated 6-1/2 ours f ight ime, n ddition o 6 hours f round esting.

Lift haring enween ain otor nd wing as investigated over road ange f he light nvelope. It as etermined that aximum peed erformance equired ower ollective itch settings ith he esultant nloading f he ain otor. Rotor

lift oading was educed o inumua? f 1

» percent f he gross eight in evel light

Flying quality nd performance ata were btained hrough comprehensive nstrumentation f he ircraft nd y xternal photo overage rom he round nd rom hase lane.

Throughout he ntire est peed ange, and n ll ested flight egimes, he KiH-lA roved o ave uitable controlla- bility ith he wo xceptions oted elow, oth f which could e liminated y ncorporating onventional ileron oll control. The ircraft ossesses igh egree f ositive y- namic tability bout ll xes t orward light peeds bove

35 nots. Positive verall ongitudinal tatic tability s evident rom pproximately 0 nots o aximum peed ttained. The lateral yclic ontrol as ntentionally ade ensitive because t as nticipated hat, ithout ailerons, t ould e needed n igh-speed nloaded light. At peeds bove 150 knots nd ollective itch ettings f 3 egrees r ess, the ateral ontrol esponse ecreased ecause f he low otcr loading. Compensation or his decrease n ontrol esponse could e btained y connecting he xisting ilerons (flaper- ons) to he yclic tick. In overing, although he ilot at no ime ad ny ifficulty ith he ateral ontrol ecause f

ill

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this yclic igging. t was ound o e ore ensitive han permitted y IL-H-8501A, Ailerons ould permit he ateral cyclic o e educed o he llowable ensitivity.

The ing-tall method of counteracting main otor orque by eflecting he irflow at he all propeller duct proved o be ntirely dequate n axiing, n hovering, nd n ideward, backward, nd ransition light.

Plight andling qualities f he odel 6H-1A during TO L tests were onsidered y he pilot o e imilar o hose f conventional ixed-wing aircraft.

round un peeds s igh

as 5 knots were experienced during akeoff with o endency to werve. orward peed andings ith partial ower proved equally onventional nd controllable*

The maximum rue irspeed emonstrated was 95 nots, which was chieved n 10-degree ive. One major actor which imited extension of he peed envelope uring his flight est program was he nadequacy f he nlet ir duct to he uried engine nstallation. The ir nlet sed or these ests recluded ull engine orsepower ue o ir star- vation at igh powers.

Vibration was not imiting actor p o he a:.iinum speeds ttained n exploring he performance id lying qualities f he 6H-1A. A hree-per-revolution ibration ospo- nent imited operation n he arlier tage f he rogram. However, tiffening of he ontrol ystem alleviated his condition.

■

All he easured tresses n he ynamic omponents er- mit nlimited atigue ife. Bearings nd other wear-limited components ere perating at oads hich would permit at east ^20 hours f ervice t he highest peeds ested.

1

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—"

■

.

F O R E W O R D

This eport unm arlzes he esults f n exploratory flight est esearch rogram o eterm ine igh peed aspects of otary-wing aircraft tilising he 6H-1A haft riven, compound helicopter. The rogram as onducted y iasecki Aircraft orporation nder S A AV L A B S ontract A -177-AM C- 15*(T).

The ehicle escribed herein, he 6H-1A, s n utgrowth of he 6H-1, reviously uilt nd ested y he ontractor n a om pany-funded rogram.

All lights er e onducted n ccordance with A A Experi- m ental Certificate f Airworthiness o. 616H dated 0 Septem ber 1955, nd ook place etween November 965 nd July I§c6,

1

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■■■■ — — ■ . —

f

CONT E NT S

SUMMARY

l l

\ F ORE WORD

LIST O P I L L U S T R AT I O N S

l l l

• | L I S T O F TABL E S

X i l

L IS T O F S YMBOL S

xl v

INT RODUCT ION

DE S CRIP T ION OF T E S T ART ICL E

TEST INS T RUME NTAT ION

2

DE S CRIP T ION O F F L I G H T T E S T S 7

P E RF ORMANCE AN D POWER DISTRIBUTION 1

VIBRAT ION

0 1

F LYING AND HANDL ING Q U A L I T I E S ANDM A N E U V E R A B I L I T Y . . 1 9

S T RE S S E S AND L OADS

9 4

TECHNICAL P ROBL E MS 7 9

CONCL US IONS

8 8

RE COMME NDAT IONS 9 0

REFERENCES

9 1

APPENDIXES ...

^ I . D E S C R I P T I O N AND R E S U LT S O F GROUND T E S T S ... 9 2 I I . P E RT INE NT DATA O N INS T RUME NTAT ION, A I R C R A F T

D R A G , W E I G H T, A ND CENTER O F G R AV I T Y ... 9 8 I I I . S TABIL IT Y AND CONT ROL ANALYS E S 2 2

DIS T RIBUT ION 6 9

vl i

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& * *xi*&**>m* MBi miHmji mL UfmW ' ** -. MW II»Xl,BlS3 U9JMmKSIii liSf9lS ■

LIST P LLUSTRATIONS

PIQIIRE

AGE

1 6K-1 n Plight

2

6H-1A athfinder

3

hree-view rawing f 6H-1A, howing C. . Envelope

k

6H-1A nstrumentation Locations ..... 3

5 6U-1A Altitude lanning Chart . 8

6 earward Flight 0 nots

2

7 6H-1A Level Flight Airspeed Calibration 9

8

echanical Efficiency f Drive ystem versus ombined Rotor nd Propeller Power or given Ratios f Rotor ower to Combined ower ...... 3

9

over ower versus Gross Weight, in round

Effect

2

10 over ower versus Gross Weight, ut f Ground Effect 3

11 otor Hover Performance. versus „ . . k P

12

ail ropeller erformance n Hover nd Vertical Climb-Propeller ower o- effecient ersus otor ower Coef- ficient 5

13

overing ropeller ower nd itch

versus udder Deflection ....... 6

11 uel low ersus Gross eight n Hover, n Ground Effect nd ut f Ground Effect. 7

15

ower versus Airspeed, ideward Flight . 0

16

udder osition ersus peed, ideward Flight

1

17

ower ersus Airspeed, earward light . . 3

vili

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: 1 8

\19

20

2 1

%

2 2 2 3

2 4

2 5

26

27

2 8

PIOÜRE

AQE

Rudder Pos i t ion versus Speed , Rearward Pl ight I 1

Rate of Cl imb versus Speed 5

Rotor /Wing Lif t D i s t r i b u t i o n in Cl imb . ... 7

Autoro ta t ive Rate of Descent versus Speed

8

Rotor /Wing Li f t Dis t r ibu t ion in A u to r o t a t i o n

9

Autoro ta t iona l Ent ry at 48.5 Kn o t s 0

A u t o r o t a t i o n Ent ry at 139 Kn o t s l

Vert ical Cl imb; R o to r F igure o f Meri t versus Nondlmens lona l C l i m b Veloc i ty

3

Vert ical Rate of C l i m b ve rsus Power

4

16H-1A Pl igh t Tr im Parameters

5

Rotor Power Coeff i c i en t versus Advance Ratio. 7

29 rope l l e r Power Coeff i c i en t v e rsus Advance Ratio .

3

30 ower Versus Speed , Forward Pl igh t o

31 urb ine Power versus Speed , Forward Fl igh t

5

32 nve lope of Mi n i m u m Power versus

Speed

7

33 i v i s i o n of Power Between Rotor and Prope l l e r versus Airspeed . ... 9

34 otor /Wing Lif t D i s t r i b u t i o n versus Speed.

Q

35 uel Flow versus Horsepower

3

Ix

' m m

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FIGORE

AGE

36 udde r Pos i t i on ve rsus Speed, Level Fl igh t

9 4

37 round Rol l Dis t ance ve rsus Li f t -Off Speed

100

38 st Haraon ic Ver t i ca l Acce le ra t ion at Pilo t Sta t ion ve rsus True Airspeed 10 2

39 nd Harmon ic Ver t i ca l Acce le ra t i on at Pi lo t S t a t i o n ve rsus True A i r speed 10 3

40 rd Harmonic Ver t i ca l Acce le ra t i on at Pi lo t Sta t ion ve rsus True Airspeed

104

4 1 t h Harmon ic Ver t i ca l Ac c e l e r a t i o n at Pilo t Sta t ion ve rsus True Airspeed ...... 10 5

4 2 t h Harmon ic Ver t i ca l Acce le ra t ion at Pilo t S t a t i o n ve rsus True Airspeed 10 6

4 3 th Harmonic Ver t i ca l Ac c e l e r a t i o n at Pilo t Sta t ion ve rsus Tzue Airspeed 107

4 4 t h Ha rmon ic Ver t i ca l Acce le ra t ion at Pilo t Sta t ion ve rsus True Airspeed 108

45 t h Harmonic: Ver t i ca l Acce le ra t ion at Pilo t

Sta t ion ve rsus True Airspeed 109

4 6 th Harmonic Ver t i ca l Acce le ra t i on at Pi lo t Sta t ion versus True Airspeed 110

4 7 0 t h Harmon ic Ver t i ca l Acce le ra t i on at Pilo t Sta t ion ve rsus True Airspeed I l l

4 8 otor-Transmiss ion Main Case Vibra t ion 113

i

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FIGURE PAGE k9 all-Rotor rive haft earing-Support

Vib ra t io n

i l *

50 er t ica l Vibra t ion , Pro p e l l e r S haf t 115

5 1 ongi tud ina l Vibra t ion , Pro p e l l e r Shaf t 116

52 ateral Vibra t ion , Prope l l e r Shaf t .... 117

5 3 er t ica l Vibra t ion , Fuselage, S ta t ion 9 .5

(Loca t ion of Instrument Panel )

118

54 ongitudinal tability-Level light, Stick osition ersus peed 20

55 ongitudinal tick osition hange ersus Elevator Trim: 130+ nots 21

56

ontrol osition ersus irspeed, Rear- ward light

22

57

itch esponse-Rearward light, itch Down

23

5S im e istory of ontrol Motions uring Steady over

, 25

59 irplane ode cceleration, orward . G., 26

60 uick top, Forward . G

27

61

irplane ode cceleration, Aft C. G. . ,

28

62

uick top, ft CO

29

63

elicopter Mode cceleration 50

6'» uick top

31

65 itch esponse-Hover, Pitch own, C. G. 8.5

33

66 itch esponse-Hover, Pitch p, C. G. 8.5

34

67

i tch Response-Hover, P i t c h Down, C . G . 5 .1

1 35

x l

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?

*

FIQURE

AQE

166 86 Roll esponse-Left ideward Plight, Roll

Right

87 itch esponse-Left ideward light,

Pitch own

67

88

itch esponse-Left ideward light, Pitch p

68

89

aw esponse-Left ideward light, aw

Left

69

90 aw esponse-Left ideward light, aw Right

70

91 oll esponse-Right ideward light, oll Left

71

92

oll esponse-Right ideward light, oll Right

72

93

itch esponse-Right ideward light,

Pitch own

73

94 itch esponse-Right ideward light, Pitch p

7*

95 aw esponse-Right ideward light, aw Left

75

96

aw esponse-Right ideward light, aw Right t

76

97

oll esponse n over-Acceleration ersus Lateral tick osition 77

98 ynamic tability, oll eft, peed 50 Knots, C. . 9.7

78

99 ynamic tability, oll ight peed 50

Knots, C. . 9.7

79

*

00

overing Turn-Left 82

10 1

overing Turn-Right 83

10 2

ontrol osition ersus irspeed,

Autorotation 85

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T

FIGURE

A G E

103 i r ec t iona l Stab i l i ty, Late ra l St ick and Rudder Pedal Pos i t ion versus Sid e s l i p Angle

1 86

10 4

ontrol osition ersus ate f limb (Vertical)

88

10 5

ateral ontrol osition ersus irspeed . 89

10 6

elicopter Handling Quailties-Criteria n

Roll From eference )

92

10 7 otor Blade lapwise ending ersus irspeed, evel nd limb, teady omponent, Station » 6

98

108

otor Blade lapwise ending versus irspeed, evel nd limb, lternating Conponent, tation 6 99

10 9

otor Blade lapwise ending ersus irspeed evel nd limb, teady omponent, tation 9.5

0 0

110 otor Blade lapwise ending versus irspeed, evel nd limb, lternating Conponent, tation 9.5 0 1

111 otor Blade lapwise ending versus irspeed, evel nd limb, teady omponent, Station 9.2

0 ?

112

otor lade lapwise ending ersus irspeed, evel nd limb, lternating Component, tation 9.2 0 3

113

otor Blade lapwise ending ersus ir-

speed, evel nd limb, teady omponent, tation 05.6

04

114 otor lade lapwise ending ersus irspeed, evel nd limb, lternating Component , S ta t i on 1 0 5 . 6 ,,«.,... 2 0 5

115

otor Blade lapwise ending ersus irspeed, evel nd limb, teady omponent, tation 24

0 6

/

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FIGURE PAGE

•

116 otor Blade lapwlse ending ersus Air- speed, evel nd limb, lternating Component, tation 24

117

otor Blade lapwlse ending ersus irspeed, evel nd limb, teady omponent, tation 32

118

otor lade lapwlse ending ersus irspeed, evel nd limb, lternating

Component, Station 32 119 otor Blade lapwlse ending versus ir-

speed, evel nd limb, teady omponent, tation 58.4

12 0 otor lade lapwlse ending ersus irspeed, evel nd limb, Alternating Component, tation 58.4

121

otor lade lapwlse ending ersus irspeed, evel nd limb, teadv onaaon- ent. tation 84.8

..

12 2 otor Blade lapwlse ending ersus irspeed, evel nd limb, lternating Component, tation 84.8

12 3 otor Blade lapwlse ending ersus irspeed, evel nd limb, teady omponent, tation 11.2 .«.....<

124

otor Blade lapwlse ending ersus irspeed, evel nd limb B lternating Component, tation 11.2 ,

12 5

otor lade lapwlse ending ersus ir-

speed, evel nd limb, teady omponent, tation 37.6 .

12 6 otor Blade lapwlse ending ersus irspeed, evel nd limb, lternating Component, tation 37.6

12 7

otcr lade lapwlse ending Moment ersus Blade tation . s

207

208

209

210

211

212

213

214

215

216

217

218

xv

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H

FIGURE

128

129

130

131

132

133

1 3 ^ 1

135

136

137

138

139

Rotor Blade Flapwlse ending Moment versus lade tation

Rotor Blade hordwise ending versus Air-

speed, Level nl Climb, teady omponent, tation 6

Rotor Blade hordwise ending versus Air- speed, Level nd Climb, Alternating

Com ponent, tation 6

Rotor Blade hordwise ending versus Air-

speed, Level nd Climb, teady omponent, tation 31.5

.

Rotor Blade hordwise ending ersus Air- speed, Level nd Climb, Alternating Com ponents, tation 31*5 • •

Rotor Blade lapwise ending versus Air- speed, Autorotation, teady om ponent, Station 6

Rotor Blade Flapwise ending versus Air-

speed, Autorotation, Alternating om - ponent, tation 6

Rotor Blade lapwise ending versus Air-

speed, Autorotation, teady om ponent, Station 9.5

Rotor Blade lapwise ending versus Air- speed, Autorotation, Alternating om - ponent, tation 9.5

Rotor Blade lapwise ending ersus Air-


Rotor lade Flapwise ending ersus Air-

speed Autorotation, Alternating om - ponent, tation 9.2

Rotor Blade Flapwise ending versus Air-


PAGE

219

220

221

222

223

224

225

226

227

228

229

230

> 4

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FIGURE PAGE 140 otor Blade Plapwlse B e n d in g versus Alr-

speed f Autoro ta t ion , A l te rna t ing Co m - ponen t , Sta t ion 105 .6 2 31

mi Roto r Blade Plapwlse B end ing versus Air- speed, Autoro ta t ion , Steady Componen t , Sta t i o n 124

2 32

142 o to r Blade Plapwlse B e n d in g versus Air- speed, Autoro ta t ion , A l te r n a t i n g Com- ponent , Sta t i o n 124

2 33

143 o t o r Blade Plapwlse B e n d in g versus Air- speed, Autoro ta t ion , Steady Componen t , Sta t ion 132

2 34

14 4 otor Blade Plapwlse Bend ing versus A i r- speed, A u t o r o t a t i o n , A l te r n a t i n g Com- ponen t , Sta t i o n 132 2 35

145 o t o r Blade Plapwlse B e n d in g versus Air- speed, A u to r o t a t i o n , Steady Comoonen t , Sta t ion 158 ,4

236

14 6 o t o r Blade Plapwlse Bend ing versus Air- speed, Autoro ta t ion , A l te r n a t i n g Co m - ponent , Sta t ion 158 .4 2 37

14 7 otor Blade Plapwlse B a n d in g versus Air- speed , Autoro ta t ion , Steady Componen t , S t a t i o n 184 .8 2 38

148 o t o r Blade Plapwlse B e n d in g versus Air- speed , Autoro ta t ion , A l te rna t ing Com- ponent , Sta t ion 184.8 2 39

149 otor Blade Plapwlse B e n d in g versus Air- speed , A u t o r o t a t i o n , Steady Componen t , Sta t ion 211 .2 ..... 240

150 o to r Blade Plapwlse Bend ing versus A irspeed , Autoro ta t ion , A l te r n a t i n g Co m - ponent , Sta t ion 211 .2 2 41

151 o t o r Blade Chordwise B end ing versus Air- speed , Autorc ta t lon , Steady C o m p o n e n t , Sta t ion 46

2 42

xvii

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PIQURE

152 otor Blade Chordwlse ending ersus irspeed, Autorotation, lternating omponent, tation 6

15 3

otor lade hordwlse ending ersus irspeed, utorotation, teady omponent, Station 31*5 ••

* • ••

ISM Rotor Blade hordwlse ending ersus irspeed, Autorotation, lternating omponent, tation 31.5

155 otor lade lapping ngle ersus irspeed

156 otor itch ink oad versus irspeed . .

157

otor Shaft orque ersus irspeed ....

158

otor haft ift ersus irspeed

15 9

eometry f otor haft

160

lternating Bending Moment ersus irspeed*

Rotor haft, ower Gage

161 lternating Bending Moment ersus irspeed- Rotor Shaft, pper Qage

.

16? Rotor haft orizontal eaction t ower Bearing ersus irspeed

163 Rotor haft orizontal eaction t pper Bearing ersus irspeed .

lö'l Rotor haft, orlsontal eaction t otor Head ersus irspeed

165 otor haft, ending Moment t otor ead versus irspeed

166 all ropeller haft orsion ersus ir-

speed

167

all ropeller ub ending ersus irspeed

168

ropeller lade lapwise tress ersus Airspeed, tation 3.2

PAGE

213

2M

* 4.

2 ^ 5 2 4 8

2 4 9

2 5 0

2 5 1

2 5 2

2 5 9

2 6 0

2 6 1

2 6 2

2 6 3 2 6 4 1

2 6 6

2 6 7 i

268

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F I G U R E

A G E . 1 69 rope l le r B lade P lapwi se Stress versus

Airspeed , S ta t i on 16.5 6 9

1 70 rope l le r bxade F lapwi se Stress versus • Airspeed

7 1 ^

171

ropeller Blade Flapwise tress versus Airspeed

72

172

ropeller Blade lapwise, teady ending

Moment ersus pan 73

173 ropeller itch ink oad ersus Airspeed 7 ^ 1

1 7 ^ 1 Longeron lternating Stress ersus irspeed 276

175

lternating tress ersus rue Airspeed Forward tructure, Lower eam, Left Side, tation 5

77

176

ypical M ain anding Gear oads

78

177

orward uselage ertical ibration . . .

85

178 robable rrors f ata ersus umber f Parameters

13

179 ongitudii.? tick osition ersus . 0. Position overing, eutral levator 23

180

ongitudinal tick osition ersus levator rim overing light

2I

181

itch esponse n over cceleration versus ongitudinal tick osition . .

26

182 itch esponse n over - itch p, C. G. 3.1

27

183 i t ch Response i n H o v e r - P i t c h Down, C , G. 3.1

32 8

1 84 aw Response i n H o v e r - Acce le r a t io n versus Rudder Pedal P os i t i on , C . G . 8 .5 ..

32 9

185 a « ? Response i n H o v e r - Y aw Left , C G. 5 .1

330

x i x

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a w

j*mHmBmm**rifi.-

PIQÜRE

A GE

186 aw Re oonse n over - aw ight, C. . *1

31

187 aw Response I n Hover - Y aw Left, C . 0 . 3.1

332

188

aw Response n over - aw Right, C . 3.1

33

189 ol l Response I n Hover - Rol l Left, C . Q . 8.5

3 31»

190 ol l Response I n Hover - Rol l Right, C . Q . 8.5

3 3 5

191 ul l Rrsponse I n Hover - Rol l Left, C . 0 . 3.1

3 3 6

192 ol l Response I n Hover - Rol l Right, C a . 3.1

337

193 ol l Response I n Hover - Rol l Left, C . a . 5 .1

338

19^ Yaw Response I n Hover - Left Yaw, C . 0 . 8.5

343

195 ew Response n over - ight aw, C . 0 . 8.5 ... 314

196 ynamic Stabi l i ty - Pitch Up, Speed 50 Knots, C . Q . 3.1

3 5 0

197 ynamic Stab i l i ty - Pi tch Down, Speed 5 0 Knots, C . Q . 3.1 •

351

198 ynamic tability oll eft, peed 50 nots, . . 8.6

54

199 ynamic tability oll ight, peed 5 0 Knots, C . 0 . 8.6 355

20 0 ynamic Stab i l i ty - Rol l Left, Soeed 50 Knots, C . Q . 3.1 3 5 6

201 ynamic Stabi l i ty - Rol l Right, Speed 5 0 Knots, C . Q . 3.1

357

xx

•

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FIGURE

202

203

201

Dynamic S t a b i l i t y - Yaw Left , Soeed 50 Knot s , C . Q. 3.1 ......

Dynamic Stab i l i ty - Yaw Right , Speed 50 Kno t s , C . G. 3.1

Dynamic S t a b i l i t y - Pi tch Up, Speed 150 Knot s , C . G. 9 .7

PAGE

361

362

365

x x i

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LIST F ABLES

TABLE

I

II

III

IV

V

V I

VII

VIII

IX

X

XI

XII

XIII

XIV

X V

XVI

Description f he est rticle

Flight nd lip umbers or ata sed n . . Mechanical fficiency urves

Flight nd lip umbers or ata sed n Performance urves

Forward elocity akeoff ata

Forward elocity/Fixed ollective-Pitch Run-on andings ....

Flight nd lip umbers or ata sed Flying nd andling Qualities .....

Rotor lade ection roperties nd Centrigugal orce .....

Flight nd lip umbers or ata sed n

Rotor lade Flapping Angle urve

Rotor haft lternating oads nd oments, Hover

Rotor haft lternating oads nd Moments,

Right ideward light

Rotor haft lternating oaus and oments, Left ideward light

Rotor haft lternating oads and Moments,

Rearward light

, . , ,

Rotor haft lternating oads nd Moments, Vertical Climb

Rotor haft lternating oads and Moments, Forward peed limb . .

Rotor haft lternating oads nd Moments, Level light

Tie-down est ower istribution chedule .

PA G E

8

3 6 - '

9 5 1

9 7

9 8

1 4 7

1 9 7

2 ^ 7

25*

,25*

2 5 5

2 5 5

2 5 6 •

2 5 7

t 2 5 8

2 9 6

xxii

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TA B L E

X V I I

X V I I I

X I X

X X

16H-1A nstrumentation

Summary of Measurement Accuracy

1 6 H - 1 A Weig h t and Cente r-o f -Grav l ty L og

Es t imated Drag Breakdown, 1 6 H - 1 A

PA G E

2 98

30 7

32 1

x x l l l

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L ^

ngine i n l e t pressure l o s s

L p rolling moment derivative with r o l l r a t e , foot-pound-seconds per radian

L y l i f t derivative with s p e e d , pounds per f o o t *

er second

La l i f t derivative with a n g l e of attack c h a n g e , pounds per radian

' L6 l i f t derivative with longitudinal stick displacement , pounds per i n c h

o r

rolling moment derivative with lateral stick displacement, foot-pounds per inch

N Mach number

MH calculated moment i n rotor s h a f t a t rotor h u b , inch-pounds

MLQ measured moment i n rotor s h a f t a t lower strain

g a g e location, inch-pounds

MQ pitching moment derivative with pitch r a t e , foot-pound-seconds per radian

Myg calculated moment i n rotor s h a f t a t upper bearing location, inch-pounds

M J J Q measured moment i n rotor s h a f t a t upper strain gage location, inch-pounds

My pitching moment derivative with velocity change (velocity stability), foot-pound- seconds per f o o t

M0 pitching moment derivative with angle o f attack change ( a n g l e of attack stability), f o o t - pounds per radian

M6 pitching moment derivative with longitudinal stick displacement, foot-pounds per i n c h

m aircraft m a s s , s l u g s

N engine output s h a f t RPM

X X V

. i^ggpwwv*—** ■

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(

N REP

N 6

n

n o

n t

OAT

P

P E

PSHP

*V P2 P t2 / P S6

Qp

QR

R

RSHF

%

s p

referred eng ine ou tpu t shaf t PPM

yaw moment de r iva t ive wi th yaw ra te , foo t -pound- seconds p er r ad ian

yawing moment with s ides l ip der iva t ive , foo t - pounds per radian

yaw moment der iva t ive with pedal d isp lacemen t , foo t -pounds per Inch

bl ip number ; a l so load fac to r

s t l c k - f l x e d neu t ra l poin t , f rac t ion of C

t o t a l number o f bl ips In a f l igh t

ou ts ide ai r tempera ture , degrees cen t ig rade

p e r io d of osc i l l a t ion , seconds

eng ine to rquemete r pressure , pounds per square i nch

prope l le r shaf t horsepower

ram pressu re r a t io

t o t a l engine pressu re ra t io

prope l le r torque, inch -pounds

ro to r torque, inch -pounds

r ad ius , feet ( subscr ip t P refers to propel ler ; subsc r ip t R (or none) refers to main ro tor)

ro to r shaf t horsepower

ca lcu la ted hor izon ta l r e a c t i o n at lewer bea r ing suppor t ing ro to r shaf t , pounds

ca lcu la ted hor izon ta l r eac t ion at upper bea r ing s u p p o r t i n g ro to r shaf t , pounds

Laplace var iab le

long i tud ina l shor t per iod mode t ime cons tan t , seconds

»

I

xxvl

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•

2

Tt2 / T2

V

V o

V C

V C A L

VC O R R

Vc O B S V C R E P

Vc S T D VT

V T R U E W

WR E F

W S T D W TO

f C O R R X C G

XV

X x

temperature at engine compressor In l e t , degrees cent igrade

ram temperature ra t io

l ong i t ud ina l ve loc i t y, feet per second

t r im f l igh t speed, feet per second

rate of c l imb , feet per minute

cal ibra ted airspeed, knots

corrected airspeed, knots

observed rate of c l imb , feet per m in u te

referred rate of c l imb , feet per minute

standard rate of c l imb , feet per minute

ti p speed, feet per second (sub-subscript P refers to propel ler ; sub-subscr ipt R (or none) refers to rotor)

true airspeed, knots gross weight , pounds

referred gross weight , pounds

standard gross weight , pounds

takeoff gross weight , pounds

fuel f low, pounds per hour

corrected fuel f low, pounds per hour

center-uf-gravlty locat ion , f ract ion of C

longi tudinal force d e r i v a t i v e wi th v e l o c i t y, pound-seconds per foot

l ong i t ud ina l force d e r iv a t iv e w i t h long i tud ina l s t ick disp lacement» pounds per Inch

change I n aircraft angle of attack from t r im, radians

x x v l i

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s

&SHP

AVC

6

«o

6(t)

C

n

e

e 2

M

P

P o

0

(R/C)

aircraf t s ides l ip angle , raölana

increment in horsepower due t o ver t ica l speed

increment In ra te of c l lmo due to horsepower d isc repancy

c o n t r o l p o s i t i o n ( s t ick or peda l ) . Inches

ambient pressu re ra t io

s t i ck or pedal d isp lacemen t f rom t r im input t ime func t ion , inches

c r i t i ca l damping r a t io

overal l eff ic iency

r o to r t r an smis s ion loss

change i n a i rc ra f t p i tch ang le f rom t r im, rad ians

t empera tu re ra t io at eng ine compresso r in le t

r o to r t ip speed ra t io

ambient dens i ty

mass a i r dens i ty at sea level - - ' • b - s e c ft H

dens i ty r a t io

ro l l ang le

yaw ang le

undamped na tura l f requency, rad ians per second

t

xxvl i l

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«

I N T R O D U C T I O N

T h i s report presents the f l igh t test resu l t s of the Plaseck l 1 6 H - 1 A , shaf t -dr iven ompound hel icopter. hi s program w as conducted under Contract D A M-iyT-AMC-lSMT) awarded

* by th e Qoveminen t i n M ay 1 9 6 ^ to obtain test data r e l a t i ve to the high-speed aspects of ro tary-wing ai rcraf t .

T he m a i n ob jec t ive of th e f l i gh t test program was to i n- ves t iga te the f l y i n g qua l i t i e s and performance of the 1 6 H - 1 A over th e speed range of hover to "max. secondary object ive was to g a i n a bet ter unders tanding of the interdependent de-

s ign parameters

n the compound hel icopter, such as the prope l l e r /ma in rotor power d i s t r i bu t i on and th e d i s t r i bu t i on of l i f t between the w i n g and m a i n rotor.

D E S I G N B A C K G R O U N D

In rder o ccomplish he bjectives, it as ecessary o modify he xisting laseckl 6H-1 ompound Helicopter see Figure ) to he *odel esignated 6H-1A. Principal modifications equired or he igh-speed egime ncluded: nstallation f General lectric 58 urbine; ncorporation f H-21 otor and ontrols; esign, fabrication, and nstallation of ew ransmission ystem; engthening f he uselage;

development f ew ropeller; nd eef-up f arious tructural omponents.

F o l l o w i n g mod i f i ca ti ons , the 1 6 H - 1 A completed 66 hours of ground t es t ing that i nc luded th e fo l lowing :

1 . omponent Tes t i ng a. ai l P rope l l e r and Ta i l Rudders b. ransfer and M a i n Transmis s ions

2 . i e ' D o w n Tes t in g a• i r f rame Shake Tes t b. i ng P roof Load c• ontrol Proof Load d. anding ear tress e. est ummary nd esults f. ear-Down nspection

In fo rmat ion on th e ground test phase of the contract i s g i v en i n Append ix I .

F L I G H T T E S T I N G

F l i g h t t es t ing was conducted from 13 November 1 9 6 5 to 6 Ju ly 1966 under Federal Av i a t i o n Agency Exper imen ta l A i r -

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I ^ S I5^SS* ?| HW ff* - «ffl

I

O u bß

aMMWWHil m w i illMMil illMinWliKimnilHiiimni'i

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worth iness C er t i f i c a t e N6 1 6 H da ted 10 S e p t e mb e r 1965, in accordance with the objec t ives set fo r th in t he Flight Test Agenda, PiAC Report 16-Y-13. he tes t phases were as fo l lows :

1 . ax i 2. overing n nd ut f round Effect 3* Sideward nd ackward light 1. ransition rom over o orward light

and Back o over 5 . orward Speed Cl imb 6 . u to ro ta t ion 7 . er t ica l C l imb 8 .

evel Fl igh t

9 . aneuverab i l i ty 10. ynamic S tab i l i t y 11 . TOL Performance

In th is repor t each phase of t e s t i n g is d i scussed i n i n - dependent sec t ions tha t inc lude descr ip t ion of t e s t s , tes t da ta and resu l t s , and ana lys is of resu l ts .

The f l igh t condi t ions have been assembled in to one s ec - t ion and presen ted i n Tab le XIX.

Stresses and load da ta were ob ta ined f rom many of t he

tes t phases and have been presen ted in a sepa ra te sec t ion .

&»«*»;»»

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DESCRIPTION OF TEST ARTICLE

The 6H-1A s ow-wing, ingle-engine, haft-turbine- powered ompound elicopter (see igure ), A t he uselage aft nd, he ircraft s quipped with ucted usher ropeller nd with et f ertical nd horizontal ail urfaces. The ertical ail urfaces (rudder anes) laced ehind he propeller erve o ontrol he ircraft n aw, oth n overing nd n orward light egimes. They re lso sed o roduce ide, ntitorque orce n overing. The orizontal ail surface s omposed f ixed tabilizer nd rim elevator.

The ngine rives oth he ain otor nd the ropeller mechanically» through n verrunning lutch hich llows he rotor o rive he ropeller, r ice ersa, hen he ngine is lowed ow n r topped in utorotation). The ngine utput peed s ontrolled y onventional urbine overnor, and ts ower istribution etween he otor nd he ropeller depends n he ower emanded y he lade itch etting f each.

The otor ystem onsists f odel 21-C etal otor blades nd hub ssembly. It s hree-bladed, ully rticu- lated ystem, sing ension/torsion traps for he lade e-

tention.

The hree-bladed pusher propeller onsists f he lades, designed nd manufactured y iasecki ircraft Corporation, and he ub, hich s f tandard irplane manufacture, odi- fled o ccept echanical lade-pitch ontrol n ieu f he hydraulic ontrol ystem normally sed.

As est ehicle, he ircraft as quipped ith wo seats: one or he ilot n he ight ide nd ne or he flight est ngineer 1. he eft ide. The emaining portion of he cabin as sed o ouse he nstrumentation, allast, etc.

The ing onsists f he enter ortion nd f he uter panels. The enter portion s f wo-spar onstruction nd houses he ladder-type ain uel ank. The uter wing panels are f ouble kin oneycomb onstruction nd re ealed o form ntegral uel anks The ft ortion f ach uter ing panel orms ull pan laperon.

The anding ear s f onventional, ain wheel/tail e- sign nd t ses ir-oil hock bsorbers. The ain heels re attached o he enter ing panel nd etract nboard, nto wells n he nderside f he uselage, y eans f lectrically perated crew acks. Emergency peration s y ech-

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«i •o c

p

(X ,

u bO

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a o

> c w

Ü

w c

o

I

rH

o

c

C O u Q

0) H > I

0) u

xi

on

0 ) U n H

6

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1 TABLE I . DESCRIPTION OP T E S T ARTICLE " l

b e n e r a l Uni t s

l b 5 , 7 8 2 to 8 1 3 5 » j e s i g n Qross Weight

Fuel Capac i ty (41 gals , p lus 1 60 ga l s . ) l b 2 7 9 + ^09 I

Normal Crew no. 2 1

p n p t y Weight l b 1 U 1 5 j [Overall Length - b lades ex tended f t HH

lOverali Leng th - blades fo lded f t 3 7 . 3

Height f t 11 . 2

Land ing Gear Type Retractable main wheel s - nonre- t ractable t a i l i w h e e l J

{Wheel Base f t - in . 19-9 | Number o f Ro tor s (Loca t ion & Type) — O ne m a i n rotor 1

O ne pusher prop

Istructure —* Alum, a l loy semi- j monocoque and | honeycomb f

main Ro t o r

~ F u l l y ar t i cu la ted 1 (H-21) w i t h ten-

s ion tors ion j straps

Type

•Diameter f t i\H

{Blades n o . 3 (metal) |

Blade Chord i n . is

{•Various componen t s were designed {weight was l imi ted by s t r eng th of

to di f fe ren t l imi t s . est fuse lage and l a nd ing gear.

»

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4ain Roto r (Cont^d)

TABLE I . - C on t inued

Air fo i l Sec t ion

Blade Taper

Blade Twist (Root A ir f o i l to Tip Air fo i l )

Rota t iona l Axis Ti l t to F use Ax i s

Disc Area

Total Blade Area

Disc Loading (based on 6500 pounds gross weight)

formal Opera t ion Speed

Normal Tip Speed

Tall Group

ni ts

N A C A 0 0 1 2

— C ons tan t chord

deg 7.3

deg 3

sq f t 1520

sq ft 99

Ib /sq f t 4 .28

RPM 279

f t /sec 6^3

The all roup onsists f propeller nd a ing or hroud surrounding he ropeller, upported n ront y orizontal and ertical tators nd 6 n urn, upporting n levator urface nd n ssembly f udder urfaces n he ropeller llp- tream.

Ring-Tali pan

Ring urface rea Outside)

Arm o otor enterllne rom

Vertical urface enter f Pressure

Rudder rea

Elevator rea

Horizontal tabilizer rea

Tall ropeller

ft

sq t

ft

6,i|

47.5

20.5

sq f t 20 ,3

sq f t 3.7

sq f t 4.9

Diameter ft 5.5

:, . . - ."v. . - --- , t & 4 a £ ^ H > M * . u M w m W » -- ,:»^ J^^kHtm-aik-HuMimtlIM MWÜiW ti toUl(m r iMiM l

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TABLE . - ontinued-

Wing (Cont'd) Units

n deg 40 lap Deflectlo

Powerplant

— O ne E ree

Shaft urbine

T-58-8

Type

Maximum ower (takeoff 0 in) SKP 1250 ea

Level

formal ower SHP 1050 ea

Level

Fuel ype — JP-il

R P M (at output

gears) of peed ecreaser

R P M 6000

Drive ystem

to ain otor

69.89 o atio ngine

Ratio ngine to rop 7.22 o

Controls

late ontrolling blade itch hrough

links

Upper washp

rotor

pitch

Lower tainless teel ables;

pulleys; ush-pull ods;

rotor ontrol ual ervo

units

11

-> .

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TEST INSTRUMENTATION

GENERAL DESCRIPTION

Key po in t s In the dr ive sys tem d i s t r i b u t i n g the power to t * i e ro to r and to t he prope l l e r were se l ec t ed in order to measure proper ly t he power g o i n g to t he se two componen t s as wel l as t he t o t a l power deve loped by the turb ine .

In orde r t o de te rmine f l ight loads and t o Insure s t ruc - t u ra l i n t eg r i t y with safe ty marg ins in the h igh speed f l ight

reg imes , one meta l ro to r blade was In s t rumen ted . he ro to r hinge assembly was comple te ly Ins t rumented for mot ions abou t al l axes, and one p i t ch l ink force measurement gave the pi t ch- ing moment of t he blade. imi la r ly t he p i lo t con t ro l mot ions were measured in t he cockp i t and the a t t i tude of the ai rcraf t t h roughout al l f l igh t cond i t i ons was measured about al l axes. Parameters which wou ld no t vary s i g n i f i c a n t l y in shor t t ime i n t e rva l s ( less than one-ha l f second) , such as speed , RPM, a l t i t ude , f lap se t t i ng , etc . , were taken f rom t he photopane l l oca ted i n s ide t he cabin. he res t of t he i n fo rmat ion was recorded o n th ree 18-channel osc i l l og raphs . oca t ions of t he va r ious t r ansducer s and reco rde rs are shown on the per- spec t ive drawing . Figu re 4 .

The highes t qua l i ty i ns t rumenta t ion was used t h roughou t . However, whereve r r ewi r ing was necessa ry, a l a rge r gage wire was used, as the i n i t i a l l ight gage wire i n s t a l l ed was d i f f i - cult to main ta in .

Many of t he pa rame te rs were checked by photograph ic coverage f rom the ground and/or chase p lane dur ing f l igh t t r i m a t t i t u d e t es t s . i ve camera pos i t i ons were bu i l t on the ai rcraf t t o record local cond i t i ons o f opera t ion in ad- d i t i o n to the ground and chase p lane photograph ic coverage . Several of t he se cameras were high- speed gun cameras capable of 64 f rames per second. On e of them was capab le of over 1000 f rames per second. uch photograph ic coverage a ided c o r r e l a t i o n between effect and cause, espec ia l ly dur ing the vibra t ion part of the program.

A cons iderab le number of s t r a in gages had to be r ep laced in order to improve thei r readabi l i ty, and cons t an t main tenance was r equ i r ed fo r the i n s t rumenta t ion wir ing which was a ff ec ted by fa t i gue cond i t i ons .

The pr imary r ecord ing sys tem cons i s t ed of th ree o sc i l l o - graphs wi th 18 channels each, p rov id ing a t o t a l of 54 channels . The use o f th ree osc i l lographs permi t t ed r ecord ing at th ree

d i ffe ren t speeds so that the necessa ry f requency r esponse cou ld be recorded wi th a min imum of exposed footage.

12

$

HWLW HWW

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ni I M l «. i* • I I £ k i & KM uC Cik

§ 1 U S

s

III

w c o •H tJ to Üo

c o •H pC O *» c

i pn c

I

0 ) u 3

13

ijaww a » '

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In ddition o he above-mentioned utomatic ecorders, a arge mount f bserved ata as ollected y anually e- cording ockpit instrument eadings. Some f he cockpit n- struments ere duplicated ither n he hoto anel r y re- cording n ne f he hree scillographs. This redundancy provided uick ook t he aircraft ' s erformance mmediately after he light.

Three asic ypes f ransducers ere sed o rive he galvanometers nd ndicators n he 6H-1A: 350-ohm train gages, linear nd otary otentiometers, nd utosyn rans-

mitters«

In-flight loads ere measured y eans of 50-ohm, foil- type train ages onded o he oad-carrying embers. Vibra- tion nd ccelerations ere easured y eans of eismic ype strain age ccelerometers ounted n he ppropriate ortions of he ircraft. Signal-conditioning or oth he onded nd seismic train age ridges as ccomplished y eans f ine 6-channel alance anels. These nits rovide eans f alancing he train age ransducer o ero utput at ny oad within he ange f he nstrument. They lso rovide ain control o ive he esired alvanometer eflection t he maximum oad xpected n light. Provision s lso ade or a hunt-type alibration esistor o hat ensitivity f ach channel an e hecked efore nd fter ach light, nd or

correlating he light ata with he atest oad alibration. Calibration esistors re 1* film- r ire-wound ith 0- PPM. temperature coefficient.

Position ransducers ere enerally f he otentiometer type n which ariable esistor s isplaced n roportion to he otion f om e omponent n he ircraft. To acili- tate echanical nstallation, otary ype as hosen n ome cases nd inear ype n thers, ut lectrically hey re considered nterchangeable. For hose channels hat could e moved hrough ull ravel efore nd fter each light, ignal conditioning was rovided or y he iasecki ignal ondi- tioning box. This rovides a alance otentiometer which er-

mits djustment f ircuit utput o ero t he esired e- ference osition nd ixed esistors o djust circuit ensitivity nd alvanometer amping. Pre- and oatflight sensitivity checks ere erformed y ecording while oving he

aircraft component through ull ravel rom top o top.

Those otentiometers hich ould ot be conveniently run through ull ravel efore nd fter ach light o heck sensitivity ere andled n he ame anner s he train age channels. In hese cases (rate nd ttitude yros) the otentiometer was connected cross wo egs f 50-ohm ummy

» s

in

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-

bridge abricated roR ixed recision esistors ith ow

temperature oefficient f esistance. his rovided our- arm bridge hich ould e onnected nto ne f he alance panels. s n he ase f he train age ransducers, he balance anels rovided ensitivity djustment, alance d- justment and hunt resistance calibration capability for re-

* nd ostfllght sensitivity checks.

Excitation oltage for ll ransducers discussed o ar was rovided y a '4-volt lead-acid attery hich as cotn-

•

letely eparate from he aircraft lectrical ystem. This battery as apped t , 12, l8,and k volts o hat he e- sired ensitivity f ach hannel as ithin he ange f he

gain djustment n he alance anels. The attery as n-

grounded nd as ot charged n light, ut as charged n normal attery charger t ight nd einstalled n ully- charged ondition ach orning. The voltage of he attery was onitored n he scillographs n light y mpressing the 4-volt utput on he rotor ne-per-revolution races of Number nd umber scillographs, hus ermitting he e- cording f attery oltage ith o oss n ata hannels. his voltage-monitor ystem as a ackup or he hunt alibration, since ny hange n attery oltage as eflected n he alvanometer eflection uring he hunt calibration nd ull throw ecordings hat ere aken efore nd fter ach light.

The bove entioned otentiometer ystems re nsuitable for ertain arameters hich ave epeated yclic otion nd thus end o ear t apid ate, articularly n he ase f the otor-blade osition ransmitters (pitch, ead-lag, nd flap ngles}« In hese hannels, autosyn ransmitters ere used, ince hey ave irtually o ear nd ave xcellent e- peatability. Excitation ower as rovided y static-type Inverter hat converts 28 volts .C. to 15 olts, AüO-cpa, A.C., a oltage regulation ccuracy of 1% , The output of ne hase of he autosyn s rectified, filtered, and ttenuated n he signal-conditioning ox o rovide a .C, voltage o he galvanometer roportional o ransmitter ngle. This same ystem was used or ngle of ttack, since t rovided inimum rag

on he vane and herefore minimum ysteresis.

D I S C U S S I O N O F A C C U R A C Y

A l l performance parameters (torque of rotor, t a l l pro- • pel ler, and engine, al l tachometers, airspeed, al t i tude and

outs ide- ai r temperature together w i t h th e combined values of turbine, rotor and propel le r powers, true ai rspeed, and den- s i ty al t i tude) had errors of 5 percent or less . n add i t i on , 8 ? percent of th e r emain ing data submi t ted has p robable i n- st rumentat ion errors less than percent.

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T he fact that the probable error for each channel i s usu a l ly based on maximum v a lue presented resul ts i n a h igh percen tage number for the error, s ince f l igh t data were a lways s ign i f i c a n t ly lower than the c r i t ica l value. or example , the rotor p i t ch l i n k force data have an abso lu te max i mu m error of only 4 3 . 5 pounds; but s ince pi tch l i n k loads i n f l igh t never exceeded S ^S pounds, th e convers ion to a percentage resul ts i n a mis lead ing , h igh f igure for the error. or th is reason the abso lu te errors should be cons ide red more s ign i f ican t than the percentage errors.

A s the data were reduced and plot ted, i t became apparent

that ce r ta in channe ls were p ro v i d i n g Inconsis tent data and were not as accura te as desired. hese channels were mo d i f i ed to inc rease the i r accuracy dur ing th e tests; as a resul t , the ove r a l l accuracy of the sys tem w as Improved throughout the program. or instance , the or ig ina l method of checking sensi- t i v i t y of the at t i tude and rate gyros proved to be Inadequate. These c i rcu i t s were redes igned to p r ov ide for sens i t iv i ty checks before and after each f l ight .

T he w i n g l i f t c i r cu i t s were par t i cu la r ly t roublesome from the start. round tests at v a ry i n g ambient tempera tures i n- dicated la rge amounts of dr i f t due to expans ion of the structure on w hic h th e gages were bonded. mo d i f i ed tempera ture

compensat ing sys tem w as i n s t a l l e d w hic h so lve d the drif t problem, but l i f t data i n f l igh t s t i l l proved to be inconsis tent , e v e n though ground ca l ib ra t ions appeared to be exce l len t . t i s poss ib l e that th e errors i n these channels were du e to the di f fe ren t s tress patterns exper ienced i n f l ight w i th th e wings loaded aerodynamlca l ly as compared to the concentra ted loads exper ienced dur ing ground ca l ib ra t ion . n the other hand, th e rotor l i f t measurements were determined to be consis tent , and these were used i n the dete rmina t ion of ro tor /wing l i f t dis t r ibu t ion .

I n other cases, large errors were uncovered dur ing a normal reca l lb ra t ion when the f a i l u r e to repeat the previous

ca l ib ra t ion w i t h i n the des i red accuracy ind ica ted a transducer fa i lure . n these cases the t ransducer was replaced before th e next f l ight .

S p e c i f i c accuracies for each channel are Ind ica ted i n the data presented i n A ppe nd ix I I .

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DESCRIPTION F FLIGHT ESTS

Prior o light esting he asic ircraft as eighed, and he mpty eight as etermined t 15 ounds. Gross weight as hen alculated or ach lighty and xact eights were alculated or rew, fuej and allast.

Instrumentation as et p o hat he scillographs e» corded ontrol ositions, fuselage angle f ttack, ttitude, angular cceleration, nd oads and tresses from ritical static nd ynamic ocations n he ircraft. They lso e- corded he ead-lag ngle f he nstrumented otor lade, its

pitch ngle nd ts zimuth osition, ccelerations n he aircraft,and he landing ear oads. Blade otions, lade stresses, otor nd propeller orque,and ertical rive haft stresses ere ecorded or ll lights; ther ata were e- corded nly s eeded o eet he bjectives f he pecific tests.

The hoto anel ecorded irspeed, ltitude, ime, as generator PH, ower-turbine PM , ngine orque, ir emperature, fuel low, ertical ate f llmb >a nd ideslip ngle.

Before nd fter ach light est,an nstrumentation

sensitivity heck

as onducted* This rocedure, opularly

called R a i , as ccomplished y he light est nstrumentation roup ?ter refllght nspection nd ervicing (prior o he light) and efore ostflight nspection nd servicing after he light).

All axi, over, ideward nd ackward light, nd ertical limb ests ere onducted ver he concrete at reas adjacent o he iasecki acilities t hiladelphia nternational irport, The emainder f he ests ere enerally conducted within utorotation istance f he irport.

U se f he ross eight/density ltitude atio ystem

required ata o e athered t pecific ltitudes hich re dependent pon akeoff eight, the mbient emperature,and the ross eight o hich he ata were o e eferred. Since conditions ight ary etween lights, a raph (Figure was constructed similar o he xample iven n eference ) for the eight ange f he 16H-1A hat nabled he ilot o e- termine he roper ltitude or ny est. Entering he raph with he 6H-1A ctual ross eight (less stimated uel sed before eaching est ltitude), he roper ressure ltitude can e ound or he equired eferred eight. Flights ere made t ross eights anging rom 689 o 50 0 ounds, nd at ombinations f ltitude nd emperature hich roduced

weight/density atios ranging rom 3 70 o 0^10 ounds,

17

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The vu lne rab i l i ty of the p r o p e l l e r and duct to fo re ign object damage was m o n i to r e d th roughou t the tes t program. W h i l e t es t s i n v o lv in g de l ibe ra te i n j e c t i o n of var ious kinds and s izes of mate r ia l s were not made, t he amount an d degree of foreign object damage were con t inua l ly observed under normal f l igh t opera t ions . Except for an emergency power-off l and ing in a rough f i e ld wi th t a l l reeds (see Technica l Prob lems ) , only occas iona l minor nicks and abras ions were found; these never in te r fe red wi th schedu led f l igh t opera t ions ,

TAXI

This hase f esting was evised o etermine he ability f he 16H-1A o erform axi aneuvers ith nd ithout rakes n arying ind onditions nd elocities p o 5 knots. The ollowing maneuvers ere stablished s he axi requirements:

1, Taxi pwind, ross wind, nd downwind without brakes at ground peeds p o 0 nots.

2, xecute igure ights ithout he se f rakes,

3, etermine ain otor lade o uct clearance

with C0 J rotor P M nd inimum ollective pitch hile axiing ownwind n 5-knot inds.

^, Demonstrate he bility o top he ircraft at ny ime hile erforming hese aneuvers by he pplication f rakes nd/or yclic control.

H O V E R

These ests over he nvestigation f erformance, flying qualities, nd tresses ncountered n he 6H-1A

during hover, oth n round ffect nd ut of round ffect.

Total ower equired nd ower i mined s a unction f ross eight, required as determined s a unction gross eight. Data ere recorded when bilized ith he desired otor P M nd in alm winds. Data ere athered t ncnconsecutive times for ach eight ( out f round ffect) over ange f from 70 0 o 50 0 ounds.

stribution ere eter- Tail ropeller ower of ane eflection nd

the ircraft as ta- tail ropeller itch

least hree eparate, in round ffect and actual ross eights

Flying ualities ere investigated y stablishing on-

trol ositions and esponse rates hen aking 60-degree

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fSISS%- ^*? '**>* k= :

hover ing t u rns i n winds of 0 to 20 knot s (gus t lng to 3 0 knot s ) . I n c lu d in g t u rns In winds in g r o u n d effect at the most cr i t i ca l cen te r o f gravi ty. Sta t ic s t a b i l i t y and con t ro l mot ions were inves t iga ted whi le hover ing in winds o f less than 3 knots . Var ia t ions in c e n t e r- o f - g r a v i t y loca t ion of 9.3 inches were inves t iga ted , and con t ro l p o s i t i o n s were de te rmined as r e la ted to g r o s s weight , hor izon ta l t r i m se t t ing , and t a l l prope l le r pi tch . i me h is to r ie s of con t ro l p o s i t i o n s and angu la r acce le ra t ions were obta ined about al l three axes Ind iv idua l ly for abrupt p u l se Inpu t s of pi tch , ro l l , and yaw. ests for re sponse ra tes were conduc ted f rom a s tab i l i zed hover with the most c r i t i ca l combina t ion of rudde r def lec t ion and prope l le r pi tch . he most c r i t i ca l c o m b in a t i o n was se lec ted such that a 1- inch l e f t -peda l d isp lacemen t f rom t r im (to turn to the l e f t ) brought t he pedal to fu l l t rave l .

In orde r to use cons i s ten t heights for ln -ground~effec t and out -of -ground-effec t hover ing , a ca l ib ra ted weigh ted l ine of 6 fee t and 60 feet (as measured f rom t he main l and ing gear wheels) as used fo r the re spec t ive tes t categor ies . hi s l ine was on a reel and was lowered only a f te r the a i r c ra f t was ai rborne. g r o u n d observer s ta t ioned to one s ide o f the a i rc ra f t would convey t o the pi lo t the exac t height o f his hover, a l lowing a to le rance o f 1 to 2 feet. he hover t es t s

were al l conduc ted with the a i rc ra f t f ac ing In to the wind.

In o b ta in in g a da ta poin t , the pi lo t would hover and ad- Jus t h i s co l lec t ive p i tch and cycl ic s t i ck se t t ings unt i l the a i rc ra f t was stable ; then the f l igh t eng inee r would t ake a bl ip of at l eas t 3 seconds ' dura t ion . he p i l o t would then change the a t t i tude of t he a i r c ra f t by means of e leva to r t r im and s t ab i l i ze ; then the f l igh t engineer wou ld take ano the r bl ip .

The ducted pusher-prope l le r def lec ted -vane ar rangement in t he ta i l fo r a n t l t o r q u e yaw con t ro l presents an approach that combines the func t ions of the conven t iona l he l icop te r t a i l ro to r and p r o p u l s io n propel ler. o prov ide an t l to rque force and d i rec t iona l con t ro l , the prope l le r and the vanes work toge ther. hat i s, d i rec t iona l con t ro l in hover can be main- t a ined by keep ing a re la t ive ly high prope l le r pi t ch se t t ing and a modera te vane d e f l e c t i o n angle , or by keep ing a h ighe r vane ang le but lower prope l le r p i tch se t t ing . Tests were conduc ted to inves t iga te both extremes.

The effect of the hor izon ta l t r im pos i t ion and prope l le r pi t ch was measured i n hover ing. he t r immab le elevator, because of i ts p o s i t i o n beh ind the duc ted propel ler, is effec- t ive in al l regimes of f l igh t , inc lud ing hover, and can be used

to adjust for nose -up or rose -down a i r c ra f t a t t i tudes . This

*

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w as valuated s o ts ffect n otal ower equirements nd allowable center-of-gravlty ravel.

S I D E WA R D A N D B A C K WA R D F L I G H T T E S T S

« * These tests cover the In v es t ig a t io n of performance, f l y -

i ng qua l i t ie s , and stresses encountered i n the 16H-1A ai rcraf t dur ing s ideward and backward f l i g h t at speeds up to 30 Knots.

Total power requi red and power d i s t r i bu t i on both i n and out of ground effect were recorded i n lOknot increments up

to 30 knots, i n s ideward f l i gh t , i n ca lm air.

he same data were recorded for backward f l igh t .

T he f l y ing qua l i t i e s were i n v es t ig a t ed by de te rmin ing control pos i t ions as a funct ion of speed i n both s ideward and backward f l igh t . Response rates and the effect of rap id con- t rol Inputs were a l so evaluated. he control power a v a i l a b l e was determined for each control at 30 knots i n each direct ion. T he f l y ing qua l i t i e s were determined w i t h the ai rcraft gross weigh t v a r y i n g from 6 0 0 0 to 6 3 8 2 pounds and w i t h the center of grav i ty 3 , 1 , 5 , 1 , and 1 2 , 1 inches forward of the rotor ax is .

S ideward and backward f l i g h t tests of th e aircraft were al l performed at the P i aseck i ramp area, pace car w i t h a ca l i b ra t ed speedometer was used to es t ab l i sh speed, wind- meter, suspended outside th e car dur ing the run, was used to v e r i fy speeds and to check w i n d correct ion. he pace car accelera ted to , and held , the des i r ed speed. hen the appro- priate speed had been reached, a card was d i sp l ayed f rom the car n o t i fy in g the p i l o t and f l i g h t test engineer to main t a in this speed and to take a bl ip .

From a hover, th e ai rcraft was rotated 90 ° to the w i n d and was f lown l a t e ra l ly i n to the wind. Afte r be ing s t ab i l i zed long enough for a b l i p to be taken by the f l i g h t engineer, the ai rcraf t could be rotated i n to backward f l igh t and s t ab i l i z ed for another bl ip . he f o l l o w i n g excerpt from Fl i g h t 2 6 3 Tes t Report i nd ica tes one p i l o t t echnique used i n gather ing data for th i s por t ion of the program,

" , , , T h e th i rd maneuver was r ight s ideward f l igh t of 2 0 knots, wi th fu l l lef t rudder and then va ry ing t a i l prop pi tch to br ing the ai rcraf t around to the pos i t i on 1 8 0 ° out of the w i n d and cont inuing at 2 0 knots rearward f l igh t n « «,,

The rocedure as repeated ntil he required ata oints had een chieved nd erified s eeded.

Figure shows the 16H-1A erforming earward light.

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TRANSITION ESTS

Transition epresents he flight regimes of cceleration from hover o orward light nd eceleration from orward flight ack o over.

The flying qualities f he ircraft uring ransition were stablished y eterrriniag ne ositions f ll controls (longitudinal nd ateral yclic, collective, rudder eaal,

• ongitudinal rim ane,and ail ropeller pitch) as he aircraft accelerated nd ecelerated hrough he 0 o 0-knot speed ange.

Level flight accelerations an d quick tops ere erformed etween he airspeeds f o 0 nots at center-of- gravity locations ^.8 nches and 2.M Inches forward f he rotor xis.

The following erbal ime istory describes an cceleration from hover hrough ransition nd ack o over sing technique here he ail ropeller rovides he required thrust or ranslation. Acceleration n his anner esults in uch maller pitch change nd omewhat slower cceleration hem ould ccur sing he rotor o rovide he ft com-

ponent of hrust. The echnique sed as:

1. over ith xcessive ail ropeller itch and old ft longitudinal yclic.

2. tart oscillograph ecordings.

3 « dd t a i l propel le r pi tch w h i l e m ain ta in in g at t i tude w i t h l ong i t ud ina l c y c l i c and d i rec t i on wi th rudder.

4 « ecrease co l l ec t i ve p i tch as necessary w h i l e adding t a i l propel le r pi tch (w i t h i n torque l imi t s ) .

5 . t 70 knots forward speed, decrease t a i l pro- p e l l e r pi tch to minimum.

6 , ecrease co l l ec t i ve pi tch to minimum.

7 . lare wi th af t cyc l i c w h i l e m ain ta in in g constant al t i tude.

8 , educe turbine speed control to keep the engine R P M constant (required du e to 8$

engine gove aor droop).

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9« Add n ecessa ry ta l l prope l l e r pi tch for hover.

10 . tabilize over.

11. urn off oscillographs.

The rocedure sed when aximum acceleration s esired and attitude hange n nimportant ollows:

1. over using high ane eflection with low propeller pitch etting.

2. tart scillograph ecordings.

3. pply forward cyclic nd ollective s required or orward otion,

4. ncrease ropeller pitch nd radually ower collective, smoothly ringing he ircraft o a evel light ttitude t bout 0 nots.

5. low-up nd lare re s reviously escribed.

F O RWA R D - S P E E D C L I M B These ests cover he nvestigation f he performance,

flying qualities, and tresses ncountered n he 16H-1A uring forward-speed limb t true airspeeds etween 2 nd 'll nots.

The ates f climb ere etermined s unction f irspeed, ower istribution, nd lap etting. Fiate-of-climb tests ere onducted etween peed ange f 2 o 41 nots in nominal 0-knot Increments at normal ated ower with nominal ross eight and enter f ravity. The saw-tooth technique was tilized or determining he erformance f he 16H-1A n orward-speed limb ests. During hese ests, the

aircraft lternately erformed tabilized limb and hen sta- bilized evel light or escent to ecord ate of limb as a function f irspeed, ower, and lap setting. Each climb as maintained hrough t east 000 eet f ltitude.

Flying qualities n climb ere stablished y determining control positions as a unction f airspeed, flap deflection, and enter f ravity. The tests ere conducted with centers of ravity f .6 nd .1 nches forward of he otor xis, with etracted laps at the ominal ross eight. Tests ere repeated with laps fully xtended t irspeeds f 0 o 0 knots*

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A U T O R O TAT I O N

Procedure was es tab l i shed whereby the P h i l a d e l p h i a In ter- nat ional A i rpor t runways or other adjacent clear areas were se lec ted for autoro ta t lve approaches, ormal ly, at a prede- t e rmined alti tude, a power recovery was executed, A procedure was e s t ab l i shed whereby many test f l igh t s that were conducted i n other phases of t es t ing ended w i t h an i n t en t i ona l autorotat ion.

T he f i rs t autorotat ion i n the 1 6 H - 1 A became an emergency when power could not be appl ied because of a malfunct ion i n the turbine speed se lec tor sys tem (see "Techn i ca l P roblems") , Al though this was th e f i rs t autoro ta t ion attempted w i t h t h i s ai rcraf t , th e p i l o t was able to complete a successful autoro tat ion to touch down i n to a se lec ted clear area. u r in g the f la re Just p r i o r to touchdown, th e heign t over a ground pro- t rus ion adjacent to the c l ea r i ng was misjudged and the r i ng - t a i l was dragged, shear ing the t a i l wheel and damaging the bottom of th e t r a i l i n g edge of the rudder vanes. I t i s s i g n i - ficant that the r ing- ta l l protected th e t a i l p ro p e l l e r and prevented a qu ick stoppage of the d r ive sys tem and the ser ious damage that i s common to such an occurrence.

T he p i lo t t echnique i n autorotat ion was s i m i l a r to the technique used i n convent ional hel icopters . o i l l u s t r a t e th i s , the f o l l o w i n g excerpt from the F l i g h t Report for F l i g h t 2 76 i s presented;

" , , , T h e ai rcraf t was t r immed out at 60 knots w i t h 11° ta l l prop pitch. C o l l e c t i v e was lowered and the th ro t t l e was reduced at 5 8 O O R P H . he needles sp l i t and the ai rcraf t entered autoro ta t ion wi thou t an y d i f f i cu l t y at a l l - T he t a i l prop pi tch was a t m in im u m throughout th e f l i g h t , , , . P i t c h at- t i tude control of th e ai rcraf t , power off, requires more longi tudinal cy c l i c movement than i t does w i t h power on, 6 0 knots i n autorotat ion was es tab l i shed; we then went to 80 k n o t s

and then approximate ly 90 to 1 0 0 knots. hroughout t h i s t ime th e rate of descent (Indicated) was pegged at 2 0 0 0 feet per minute ,"

" . . .Th e col iec p la te ly bottomed. R P H increase to ab o engine R P M ) and the avo id ing a di tch an clear area of what tered and th e ta i l the t ime that I tho forward airspeed an

t i v e pi tch was then conf i rmed to be com- I reduced to 60 knots w h i c h l et my rotor u t 5 9 0 0 R P H (rotor R P H equ iva l en t to 5 9 0 0

only t h ing I concentrated on was f i rs t d a cement culvert i n o rder to reach a looked l i k e grave l . A f la re was then en- wheel was felt to hi t the ground pr io r to ught i t would. e s t i l l had 30 to 4 0 knots d I had not p u l l ed an y co l l ec t i ve yet, so

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flare as continued nd ollective itch as applied nd he tall heel ounced nce r wice. W e ere ery close to ero airspeed when finally let he ront ear ome own o he ground. In he final landing, there as lenty f otor n- ertia nd he final landing f he ircraft as very smooth and ike normal anding...'.'

VERTICAL LIMB

These ests cover he Investigation f he 16U-1A erformance, flying ualities, and tresses in ertical limb.

Power equired nd otor/tail ropeller wvrer istribution ere etermined s unction f ross eig'it nd ertical ate f limb. Vertical limb ests ere concucted t 100X ain otor P M t referred ross weights of 20 0 and 0 0 0 pounds nd eferred ower arying rom GM o 226 orsepower

The ircraft lying ualities n ertical limb ere s- tablished y howing ontrol ositions as a unction f ertical ate f limb.

A tandard ilot echnique as used or ll ertical rate-

of-climb ata cquisition. From tabilized over ver he

Piaseckl amp rea, the ower was increased o redetermined level nd he ircraft climbed ertically. When he rate of climb ad tabilized, hree lips ere aken t 00-foot n- tervals.

LEVEL LIGHT

This ortion f he est rogram overs he nvestigation of he erformance^ lying ualities, nd tructural ntegrity of he 16H-1A n evel flight.

Total ower equired nd ower istribution etween he

main otor nd he tail ropeller ere determined or tabilized light by he eight/density ratio ethod entioned earlier. In rder o etermine the ptimum least power) points, power istribution t a iven peed as varied y ap- propriate changes in ropeller nd otor ollective itch. For each ombination, airspeed as aintained ith ongitudinal stick, hich lso roduced pecific attitude.

Static, irectional nd ongitudinal ontrol-fixed ta-

bility were investigated t arious ircraft peeds, ngles f attack, nd enters f ravity.

he argin f ontrol vail- able as investigated or arious centers of ravity.

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As the ata ere o e acquired sing he eight/density method, it as ecessary to ly at he roper ensity altltudo to chieve he desired eferred ross eight (w/o). The method f stablishing he roper ensity altitude and hart used or ltitude selection re discussed t the beginning f this section.

formal rocedure as to limb to he required ensity altitude, stabilize at the first speed oint, and ake a 3- o - second nstrumentation ecord. Pertinent cockpit instruments were recorded anually on he flight card s a ackup for he automatic ecorders and o rovide a uick look t the aircraft's erformance immediately after he flight. Records were aken ver rue airspeed ange from 0 o 16? nots in level light nd o 95 nots in ives up o 10 egrees. Angle of ttack as varied y eans of he levator rim ane and y using ifferent combinations f otor nd ail ropeller pitch.

Longitudinal tatic tability as investigated ver he speed ange y determining tick osition s a unction f speed n level light, climb, and utorotation. In ddition, the ircraft as trimmed t a pecific peed, and he changes in tick osition ere determined which ould roduce peed

changes from rim f 10 m ts and 20 nots.

Dynamic tability as investigated t rim peeds f , 50, nd 50 nots t wo enters f ravity (8.6 o .7 nches, and .1 nches forward f otor xis) and t ominal ross weight f 10 0 ounds, ith flaps retracted. The est rocedure as to ecord n scillographs a ime istory of he control otions and ircraft responses from which he eriod and amping hi.racteristics f he aircraft ere derived. At no ime as it ecessary for he ilot to se an bnormal control recovery rocedure, and e as reported hat the aircraft appeared o ave ery high egree of ositive dynamic tability.

It as found hat originally lanned tick-free stability tests ere ot feasible because the 16H-1A ockpit controls act hrough rreversible ydraulic ervos and re not quipped with entering ungees of ny ind. Thus the tick will tay in ny osition n hich t is laced, regardless of he force at the output end f he system. In he beginning f he rogram, the servos ere located nder he cockpit floor, and friction as so low hat the tick ould all ff n any di- rection f ts own eight. After he servos ere relocated n the ylon (see "Technical roblems"), there as nough friction o revent his falling ff, ut tick-free tability as still ithout significance.

27

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-

The hugold r onoperlod scillation cnaracterlatics were btained y olding ontrols ssentially fixed nd l- lowing he ircraft to eviate from ts trim peed nd itch attitude nd y ecording he ime istory f he nsuing onp;- period scillation.

For urposes f irspeed alibration, he osition rror affecting he ltot-static ystem was determined y he round Speed ourse Method. Several peed ourses ere laid ut adjacent o unways at he hiladelphia nternational irport. The löH-lA as flown ver hese courses of nown istance in both irections t n ltitude f ess han 00 feet so hat

transit tiroes could e ccurately measured (+.1 econd). The selection f he articular ourse as n he asis f inimum crosswind. True irspeed as determined from he esultant time/distance ata.

Figure 7 ives the results of he level light calibration f he irspeed ndicating ystem. Since he nstrument eliability below he peed f 0 nots is low, he calibration s nly useful or peeds bove his alue.

The alibration urve as onstructed n w o tages, coinciding with he peeds concurrently being ested. The first eries f ests overed n ndicated irspeed p o 29

knots, and he econd eries covered he peed ange bove his speed. For he irst eries, he urve as rawn s hown y a otted ine, ince t hat ime he oints from lights ^02 and 11 ad not et een btained. It as ubsequently orrected s hown y olid ine. It as xtrapolated s hown in hantom ines or peeds bove 60 nots (indicated) e- cause ower imitations revented alibration t igher peed by he ethod sed. Maximum eviation f easured oints from the final alibration urve s less than nots, nd verage deviation s less han 3 nots.

STOL

This ortion f he est rogram overs he nvestigation of he TO L performance f he I6h-1A nd he tresses n- countered uring imulated TO L perations.

STOL erformance as valuated y conducting orward velocity akeoffs nd andings ver easured oncrete perating trip. STOL perations ere conducted t various cross weights nd ith ifferent flap ettings. In ddition o e- cording power nd light data, landing oads n he landing gear hock trut ere easured nd ecorded. Since he 16H-1A landing ear as ot esigned or TO L weights, the ests ere actually conducted t eights for hich T O L peration ould

28

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- *

2 0 0 -

^150 o

a . C O K 10 a

Q

C Q M

5 0

OO " 1

I N D I C AT E D A I R S P E E D - K N O T S 2 0 0

F igu re 7 . 1 6 H - 1 A Lev e l P l i g h t Airspeed C al ib ra t i on .

2 9

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1

have been poss ib le , TOL opera t ion was s imula ted by main-

t a i n in g reduced co l lec t ive p i t c h se t t i n g s t h roughout the take- of f run.

Maximum s t a t i c t a l l - p r o p e l l e r horsepower with in brake h o ld in g power was set p r io r to the s t a r t of the t akeo ff run. Afte r brake re lease , the pi lo t Inc reased prope l le r pi t ch in orde r t o ob ta in max imum thrust . he a i r c ra f t was a l lowed to f ly of f withou t a t a k e o ff ro ta t ion . he co l l ec t ive p i tch s e t t i n g was he ld cons tan t du r ing the en t i re t a k e o ff run. hi s procedure did not permi t the m in im u m t a k e o ff di s t ance capab i l i ty t o be ach ieved , s ince u t i l i z a t i o n o f the fo rward componen t o f ro to r t h rus t cou ld have inc reased t he fo rward acce le ra t ion cons ide rab ly.

STOL approaches and l and ings were a l so performed wi th a cons tan t co l lec t ive p i t c h se t t ing . he prope l le r p i tch was var ied to con t ro l t he ra te of descent and was reduced to minimum jus t p r io r to touchdown. Long i tud ina l cycl ic con t ro l was used to con t ro l ai rspeed, which was main ta ined at 60 knots dur ing t he approach and was then reduced by a f la re to 43 knot s Jus t p r io r t o touchdown«

*

D ATA E D U C T I O N M E T H O D S

All erformance ata have een educed o tandard ea

level 59 egrees Fahrenheit, 9.92 nches f mercury), onstant ross weight, nd onstant otor P M . M ethods f orrection o hese tandards followed hose n Reference , nd included orrections for pressure altitude, mbient emperature, nd irspeed alibration rrors. For evel light ata, correction n ower as ade or ncidental ate f limb r descent. For onstant-power limb ata, correction or ate f climb as ade or inor ariations from tandard ower.

The pecific quations mployed n he ata eduction program re escribed elow.

For he urpose f omputing nstantaneous weight for test blip, he uel low uring ny ne light as een s- sum ed onstant from akeoff o anding. Therefore, he vehicle weight for ny lip s

W T0 [ ( n + 1 ) (--I2)] (1)

where the fuel comsumpt ion in t akeoff nd l and ing por t ions of the f l igh t ha s een accoun ted for in the propor t ion .

3 0

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The mbient ressure nd ensity ratios are given s

60 l-(6,88Hx CT6 ) ] 5 , 2 5

2a)

and

& r 288o LOAT 273J

2b)

respectively.

Equation (2a) yields the ressure ratio f he tandard atmosphere o etter han hree ignificant figures, o igher

than 0,000 feet.

Calibrated lrspeeed n orward light s obtained rom indicated irspeed s read rom he hoto anel y eans f calibration curves. True airspeed an e calculated s

TRUE ~ ^-

3)

Turbine ower s calculated irectly rom engine orquemeter pressure s ead rom hoto anel nstrumentation,

E S H P .711 (PE lW

i,)

Rotor nd ropeller haft ower re determined rom oscillograph train age nstrumentation s

^ R n,5xbj t öö6

5)

QpxN P£HP .22x63,000

6 )

where he constants 21,5 and .2? are he respective gear ratios etween ngine nd otor nd etween ngine nd ropeller.

The ollowing orrection s pplied o eight o ccount for mbient ir ensity,

u WSTD o

7)

Rotor nd ropeller ip peed nd hrust nd ower oefficients re btained rom equations 8) hrough 12), here the inal ubscripts nd efer o otor nd ropeller, respectively.

31

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r

- R t Rp N 30x21 .5

( 8 )

VT p

CTR

t Rp N

Mechanical Eff ic iency

3 0 x 2 . 2 2

WS TD

r . 50 RSHP Cp R PO«ARV

, , 50 PSHP P P PooApVTp3

( 9 )

(10)

(11)

(12)

Turb ine shaf t torque, ro to r shaf t torque, and p r o p e l l e r shaf t to rque were al l measured Independen t ly by three sepa ra te

to rquemeters . hus, ESHP, RSHP, and PSHP, computed f rom equa- t i ons ( * l ) , (5) , and (6), re spec t ive ly, represent Independen t de te rmina t ions of these quan t i t i e s . he mechan ica l losses m the dr ive s y s t e m can be asc r ibed to th ree separate sources: (1) a ccesso r ie s , such as pumps, gene ra to r, and o i l -coo le r fan, which are re la t ive ly cons tan t ; (2) losses in the t r ans - miss io n sys tem be tween tu rb ine and p r o p e l l e r shaf t , which are a func t ion of p r o p e l l e r power ;and (3) losses In the t r ans - m is s io n sys tem be tween t u rb ine and ro tor, which are a func t ion of r o to r power. unc t ions (2) and (3) are not the same, s ince d i ffe ren t gea r t r a ins are involved. ence, t he overa l l l osses depend not only upon t he t o t a l use fu l power but also upon i ts di s t r ibu t ion . Accord ing ly, p l o t s were made of e ff ic iency versus combined power ( ro to r plus prope l le r ) , fo r d i ffe ren t ra t ios of r o to r power to combined p o w e r , , f rom measuremen ts t aken dur ing the var ious types of s teady-s ta te f l igh t (hover, ver t i ca l cl imb, fo rward cl imb, level f l lgn t ) . nese p lo t s are shown i n Figure 8 : (a, b, c, d) and a susimary of al l four Is shown i n Figure 8 (e).

In t he p r e s e n t a t i o n of f ina l curves for hover, ve r t i ca l c l imb, and level f l igh t , ro to r power and prope l le r power are der ived f rom the i r re spec t ive p lo t t ed power coeff i c i en t s .

#

3 2

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J.» A

RSxiP r 75 o

1

.85

1 -

RSHP PSHP

1.0

-

- G /

-

.9 -

€ 3Ü tl

f% -

- X O > (a) -

" .8

- -

M tJ M B t i R f a U

^ i . l o M z < o

- RSHP >5 o .65 -

RSHP S HP

1.0

-

o

-

i

0

G «i. -

.9 -

U

0 ^

^G o o -

-

s 0 < )

(b) -

.8 — i— — 1— . —L— _ _i_ i, — L— —L —1—

200 300 i l 0 0 00 600 700 800 9 00 1000 1100 1$00

SUM OF PROPELLER AND ROTOR S H A F T HORSEPOWER

Figure 8 . Mechanica l Eff ic iency of Drive S ys tem versus Combined Rotor and Prope l l e r Power fo r Oiven Ratios of Rotor Power to C ombined Power.

3 3

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■

"

i

1 . 1

1 . 0

. 9

ü

U M O M

w

< Ü M

< X o w

. 8

1 - risa? RSHP PSkF "

35 c .i»5 -j

u i US

□ -1

u □ 3 Jl— -1

[- E x^ 13 □ -j

[- X 3

I i (c)

*" 1

L ;

1 . 1

1 . 0

• 9

[- 1 8SHP

i L-.15 o .25 -j R S H P ' St P

1 - -j

[- ^ —S

^A A

-1

k 4 -

l- (d) -

i L-J-,- » « i, „L. _ i 1 -

, 8 2 0 0 30 0 i lO O 0 0 0 0 0 0 0 0 90 0 1 0 0 0 1100 1 2 0 0

S U M O F P R O P E L L E R A N D R O T O R S H A F T H O R S E P O W E R

F igu re 8 . ont inued.

3

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4

«

1 . 1

1 . 0

2 0 0 30 0 0 0 5 0 0 0 0 0* 80 0 0 0 1 0 0 0 11 0 0 1 2 0 0 S U M O P P R O P E L L E R A N D I J T O R S H A F T H O R S E P O W E R

N O T E : S ee Ta b l e I I for l i s t of F l i g h t and B l i p numbers plotted i n th is figure.

Fig u re 8. -Cont inued .

3 5

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m a m m uMW- m

Turfoin« ower s erived ron he ppropriate om bination f rotor nd propeller ower, sing he efficiencies f Figure .

TABLE I I . F L I G H T AND B L I P N U M B E R S FOR D ATA U S E E I N M E C H A N I C A L E F F I C I E N C Y C U RV E S ( F I G U R E 8 )

FLIGHT B L I P S FUGHT B L I P S F L I G H T B L I P S

P5 5 5 2 9 6 2 3 5 0 4 , 6 , 8

P * 7 6 3 0 1 1 . 2 3 5 5 7 £ 5 9 2 3 0 2 4 3 5 7 1 . 2 § 6 1 5 ,8 3 0 1 2.3 3 9 9 6 hen l .B,*, 3 0 5 1 4 0 1 1.M.7

7 , 1 0 . 11 . 1 5 , 1 6 , 1 8 ,

3 0 6 5 . 6 4 0 2 2 . 3 3 0 7 2 4 0 8 5 , 6

1 9 3 0 6 1 . 6 4 0 9 2 1 b ? * 1 . 2 , 7 3 0 9 5 4 1 2 1 .2 ,3 ,5J p90 1 . 2 , 3 , 7 , 3 1 7 2 . 5 , 6 , 9 , 8

9 1 0 4 1 3 2 bn 3 3 3 2 2 ,3 4 1 6 3 , 7 1 p92 4 , 6 , 7 , 8 , 3 3 3 1 4 2 0 7 . 8 , 9

1 1 3 3 7 3 . 4 4 2 1 2 , 3

J 2 9 5 2 , i »,7 ,9 3 4 4 3 . * . 7 , 8 , 4 2 3 1 9 , 1 0 4 3 8 1 . 2 . 3

The ollowing m ethods mployed n he ata eduction program are presented with eference o he pplicable light m aneuver.

Level Flight

Having derived he ower om ponents (ESHP, SKP, SHP) as unction of gross weight nd peed, he orsepower orrection ue o n ncidental ate f l imb s

S T D ASHP (R/c) J 7 T O w

( STD) 1 0 (13)

36

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W ith he xception f hree lips t peeds n xcess f 173 nots, hich er e n ives (10 egrees ax imum) , he ean rate f limb (absolute alue) or all ther lips sed or plotting as 150 eet er minute, o hat his orrection as normally uite mall, f he rder f 5 orsepower.

* xcept or he hree lips ust mentioned, he orrected powers re

ESHP SHP ASHP,R /C)

14)

R S H P (|||P)ESHP

15)

PSHP (IÜF)E SHP

l >

For he hree xceptional blips, he eparture f he flight ath rom horizontal as onsidered ufficiently arge (up o 0 egrees) hat he ower istribution ould e igni- ficantly istorted y he orrections f equations (15) nd (16). Since he rimary ffect n hese ases as or ravity to roduce om ponent f propulsive orce long he light path, he orrections wer e applied s

A R S H P A E S H P Jn2

Y

17)

and APSHP • AESHP cos 2 T

18)

when Y - light ath ngle o orizontal

(positive ownward)

. -1

0BS " sln 1.69x60 TRÜE

19)

Power oefficients for he otor nd ropeller re hen

ulculated rom equations (11) nd 12).

In rder o present ngine uel low erformance he ollowing orrections ave een pplied o he ata.

Calibrated irspeed ecom es

V V C0RR TT

37

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f r o m which Mach number is determined;

M - _v£ Q f i E _ . - . - i»9.c i.BCöAirrnr^

2.0

The ram t empera tu re and pressu re are ca lcu la ted as

Tt y - 1 ,2H 2

22)

2

and

(23) t / 2 " (Tt2 V 1 ' 5

The verall ngine ressure atio s

? H m V 2 (1-Li) ( L-Le) (2H}

The ressure nd emperature atios t he ompressor Inlet re

« 2 «0 (I-4)

82 ( 2HF"J 6P

(25)

R A M

Specific eat orrections o haft orsepower nd uel

flow re ncluded nd rise ue o ariations n ompressor Inlet otal emperature

T2 88 2 73

The orrection actors re

Kw .001025 2 .986

and

Kp .99 .5 0"" 2

Finally e btain orreetöd uel low;

W

Wf C0RR W6 2\ / T 2

(26)

(27)

(28)

(29)

38

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and

orrected

ngine haft

ower

s;

ESilP CO R R Kp62 B^

30)

Forward l imb

Test ata eduction nvolves orrections o ate f limb due o ariations rom rogrammed ower evel. Data re re- sented t onstant ltitude ea evel tandard.

T he ncrement ue o ariation n ower evel s

33,000 (105 0-§p^)n A Vc ÜT"

31)

where 1050 epresents ea-level o r ma l ated power.

Then eferred ate f limb s

Hover

WSTD, SHP, nd SHP er e om puted y eans f quations (7)i 5)i nd 6). p nd ^ or he otor er e btained y ublng equations l)and 13) , and p as lotted gainst CT ) 3 / (Figure 1). traight ine, hich s he roper unctional relationship ccording o omentum heory, as hen itted o the ata points. he lots f otor horsepower ersus ross weight Figures nd 0) or tandard ea evel er e hen constructed rom hese C p < p plots.

The all ropeller hrust (TpR0 p) In over s directly

proportional o otor orque, nd t s asily h o w n hat he tall ropeller hrust oefficient T s roportional o he rotor ower oefficient C pR .

TPR0P • i R

Where , Is onstant etermined y ircraft eometry,

T

K ^ORSHP X R

„.T PROP

Ki —VT

33)

39

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P E R F O R M A N C E A N D P O W E R D I S T R I B U T I O N

H O V E R

Po wer requ i red i s plo t ted agains t gross weigh t I n ground effect (F igure 9 ) and out of ground effect (F igure 10) . Th ese f igures show the total turbine power requi red , as wel l as the components of rotor and propel ler power. Rotor and pro- p e l l e r power were d e r iv ed from the fa i red l i n e s of Figures 11 and 1 2 , respec t ive ly. hown for reference on Fig u re 11 are f l i gh t test resu l t s (per rotor) on an H - 2 1 B and an H - 2 1 C he l icop te r obta ined from data I n References 6 an d 3 , respec t ive ly. he power was measured I n tne tests of References 6 and 3 by reading . I n the cockpi t , m an i fo ld pressure and R P M and by ob ta in ing engine brake horsepower from engine performance curves corrected for ambient temperature and pressure. T he 1 6 H - 1 A rotor power was obta ined from torque measurements at th e rotor shaft . Co n s id e r in g these different measurement techniques, the rotor power measured on the 1 6 H 1A I s I n good agreement wi th the H-21 data (I n general w i t h i n 3 percent ) .

A n at tempt was made to f i nd trends I n power required as affected by e leva to r t r im. he poin ts show no d is t ingu i sh - able trend, however, and only a s in g le set of curves has been drawn through the points I n Figures 11 and 12 . ote that F igu re 12 (propel ler power coeff ic ien t ) shows poin ts obta ined I n ver t ica l c l i mb as we l l as hover. Sin ce the pro- p e l l e r I s opera t ing I n essen t i a l ly the same envi ronment dur ing ver t ica l c l imb , the points fa l l on the same curve, except for a bi t more scatter, wh ich I s probably at t r ibutable to th e d i f f i cu l ty of main ta in ing zero a i r speed dur ing a pro- tracted cl imb.

Pro p e l l e r power I n hover I s used to generate s ide force for antl torque and for yaw control. igure 13 shows how th e power requ i red for antl torque var i es w i t h rudder-vane angle. A s expected, the power and propel le r pi tch decrease w i t h increas ing vane def lec t ion . F or m a x i m u m h o v e r performance, p i lo t technique w as to ho ld fu l l le f t rudder pedal and make minor yaw correct ions by means of the propel ler-p i tch- control beep swi t ch located on the co l l ec t iv e p i t ch handle . F or normal operat ion, however, the technique was to use enough propel ler pi tch to prov ide at leas t a 1 - lnch marg in of lef t pedal , wh ich was equiva len t to 3 degrees of vane def lec t ion . I n a s ide wind , when power i s not a l i m i t i n g factor, more p ro p e l l e r pi tch and less le f t rudder can be used for antl torque, l eav ing mere le f t rudder a v a i l a b l e for

yaw control.

n

i-

. ■■ - - -

8/12/2019 675443


o o 00 CO

o o A

a- f CO

• 0

*

o 0 ) o «H ♦ o «-I • CO

•0 c 3

o 0 o u VD 0t- c

c /l « OQ p02 £CUD 8 1 t-O H

0*

1

0)

n ot« n OK O 000 (4VOM O u

» n 3

C O n OOT b00 0 ) 3-DC > VOO

0 (L i 00 U \o 9)

> O

0 * 0 o\ VO I tn

3

•H 0 PE I 0 * C \J •• ITi

0000

1 O O

-

D

*

* s a a H Q N n H a M o a a s n o H

42

o sr

8/12/2019 675443


r-i a s co r« - vo ir> -

s a a H d N n H HawodasaoH

^3

■■ ■ ■ i Hm m ^f mm u i •iMmUMtrntm» MHtrnttMianm tirm-im

8/12/2019 675443


—

.22 24 26 28 30 32 34 36 38 40 42 44 , 3 6 .38 40 4 2 44 .46 .48 50 52 54 .56 .5 8

CT X 102

Figure 11. Rotor Hover Performance, C p versus C I J « ,

«

4 4

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4

^0 - L NaiOL MaOO HMOd S'iigdOHd-qiVJ,

15

imtimim&am***

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—-

(dii aavia ,V ) 0 3a - HOiU H 3 11 3 d O H d

dH - H 3 M 0 d H 3 T I 3 d O H d

c 0

•H PÜ

a > rH Q

^ a >

=J t- ^3 •a

n

iC O^r S J

3 a

O t. tr\ 0 > CV £ra O

O 3 W

ao Q L . OJ

1 o a- ^S « V 0

M U

Orvj 6d O J U fe

vö W fH Q )

r-t KrH

r> i U 4 » i-s Q O.

Q O S * c c a. 00

W ) B C

_ &-. -rl ^ W f.

.J 1 F - 4 ^

5 3 c b0 1- Tl a , ft .

V

4 6

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•

s

O Oo

oo * co e** >o -H O

•«M h^«w

o > o o-o VO Ä \o 3

C O •HI U

o w a

OQ -CV. oz too =r E3 •H SQ §;*J

04 3 0 O

1 to »•o

OH Q C OK si

f\J O SVO M Pw « o s >»

« -• C O »Vi

O O ow oo H OK fccg VO O c

rH3 u o 1U

fado • o •a - CO Hirv

« U 3

bO o H o fe vo iT »

ooZT LH

unoH ad aNncd woid 3nd

17

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-

1 I I

Turb ine power In Figures 9 and 10 Is der ived f rom t he coab lned fa i red curves of r o to r and prope l le r power ( shown as do t t ed l i nes In Figures 9 and 10), cor rec ted for mechan- i ca l losses as exp la ined under Data Reduc t ion Methods .

Throughout the range of gross weigh ts tes ted , 5200 to 8800 pounds referred weight ( refer red t o sea level , 59 degrees Fahrenhe i t , and s tandard ro to r speed; actua l weight r ange was 5700 to 7500 pounds) , t a l l p r o p e l l e r power Inc reased gradual ly f rom 12 percent to 25 percent o f to ta l power required. This Is In good agreemen t with t he r esu l t s ob ta ined In aeasurements o f power and s ide force dur ing the ground t es t phase o f t he program. he balanced power requ i remen ts of t he 16H-1A con f igu ra t ion are such that the In s ta l l ed power as de te rmined by high-speed requ i remen ts a l so prov ides ample power f o r hover ing , inc lud ing the power requ i remen ts of the r ing - ta i l .

SIDEWARD AN D REARWAPX' FLIGHT

I

f

Several f ac to r s In f luenced t he sca t te r of the poin t s presen ted . s t he program progressed , p i l o t technique and ab i l i ty t o manage t he a i rc ra f t necessa r i ly improved. n an

e ffo r t to c o m p e n s a t e fo r th i s bu i l t - i n error, ce r ta in po in t s were de le ted a f t s r research showed tha t a di ff e ren t tech- n iq u e was used . o r example , al l poin t s used in r ight s ideward f l igh t were t aken while the pi lo t used the t echn ique of app ly ing fu l l left rudder, then caus ing the ai rcraf t to t r a n s l a t e by a p p ly in g cyc l i c and co l lec t ive as r equ i red whi le keep ing t he aircraf t d l rec t iona l ly or ien ted by vary ing p rop e l l e r pi tch . i t h in l imi t s , the same maneuver cou ld have been ca r r ied out by hover ing at a h ighe r p r o p e l l e r p i tch se t t ing , and then, whi le t r a n s l a t i n g s ideward, by vary ing rudde r and/or p r o p e l l e r p i tch to keep the aircraf t d i rec t ion - a l ly or ien ted .

In performing these maneuvers In ground effect , the a i r- craf t was genera l ly f lown with t he wheels 7 to 12 feet o f f the ground . he ro to r disc was thus a lmost 1/2 ro to r d iamete r f rom the g round . In hover ing , th i s would mean that approx i - mate ly 10 percen t less power would be needed to l i f t the a i r c ra f t ; but as t he a i rc ra f t begins to t rans l f e s ideward, much of th is g round effect is los t . he resul t ng d i ffe r- ence between t he po in t s t aken on f l igh t s in ground effect and out of ground effec t was so s l igh t tha t they fe l l with in the sca t te r fo r t he t es t ser ies .

As can be seen f rom the pho tographs , even the ln -g round-

effect f l i g h t s were conduc ted at an a l t i tude at which ground

effect was cons ide rab ly d imin i shed . or th i s reason, out -

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o f - g r o u n d effect da ta are d i f f i c u l t to d i ff e r e n t i a t e f r o m i n -g round-e ff ec t data.

Maneuver ing c lo se to t he ground i s not a s i t u a t i o n which lends i t s e l f to l ong bl ips of s tab i l i zed f l igh t c o n d i t i o n s , and i t was thus a source of error. Some e r ro r a l so comes f r o i a the method used to de te rmine the la tera l speed o f the ai r- craft . l t hough the best means ava i l ab le and the method spec i f i ed in the contract were used, t he da ta w o u ld have been improved i f speeds cou ld have been more accura te ly asce r ta ined .

Power versus a i r speed in s ideward f l igh t is shown in Figu re 15. l though d a t a poin t s were cons iderab ly sca t te red« the d a t a presen ted are be l ieved to ref lect the ac tua l power cond i t ion ( c o n s i d e r i n g the number o f var iab les impl i c i t in the data) . otor- and prope l le r-power curves are f a i r ed through the measured po in t s . urb ine power i s cons t ruc ted f rom the combined r o to r and p r o p e l l e r power, co r rec ted f o r mechan ica l l o sse s as exp la ined under Data Reduc t ion Methods . The poin t s fo r zero speed are t aken f rom t he hover curves . Figures 9 and 10. ess nower is r equ i r ed f o r f l y i n g to the l e f t than to t he r i gh t , because of t he inc reased t a l l prope l le r p o w e r needed t o f ly t o t he r igh t whi le def lec t ing

ai r to t he lef t . he t r end of the poin t s i nd ica tes tha t t he most power is r equ i red at about 18 knot s to t he r igh t , with power requ i red dropp ing as the speed changes f rom th i s va lue .

The r o to r p o w e r requ i red decreases wi th speed , at an i nc reas ing ra te , up to 30 knot s in each direct ion. he sm a l l di ff e rence in r o t o r power between f l igh t to the lef t and to the r igh t may be a t t r ibu tab le to the fac t that t he plane o f the r o t o r i s at a g rea te r ang le in f l i gh t to t he lef t , and/or that the a i r c ra f t is ro l l ed more in f l i gh t to the le f t t han in f l ight to the r igh t . he reason fo r the grea te r r o to r ang le in f l i gh t to t he lef t is that the r o t o r must overcome the s ide component o f th rus t f rom t he t a i l propel ler.

Tai l p r o p e l l e r power r equ i r ed increases at an i n c r e a s in g ra te f rom a m in im u m at about 20 kno t s to the lef t . t becomes l i near p a ss i n g t h rough the hover po in t and Increases at a constan t ra te to 30 knot s t o the r ight . s speed t o t he l e f t Increases , less t a i l p r o p e l l e r power is needed because i t is augmented by t he di rec t iona l s t a b i l i t y of the a i r c ra f t (see Figure 16 , showing rudder vane def lec t ion i n s ideward f l igh t ) .

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Power versus a i r speed in rearward f l igh t Is shown In Figure 17. Because o f t he inherent danger impl i c i t in f ly ing rap id ly backward c lose to t he ground , the in -g round- effec t rearward f l igh t s were f lown at an a l t i tude o f aoout 15 feet , ra the r than 5 t o 7 feet as in t he hover t es t s . Since some ground effect t ends to be los t in t r a n s l a t i o n a l f l igh t , in any case , the r e s u l t i n g da ta are such that in g ro u n d e ffec t cannot be e ffec t ive ly di ffe ren t ia ted f rom out o f g round effec t , and only one set of curves is drawn. Roto r power d?c reases with inc reas ing rearward f l igh t speed with in t he r ange tes ted , as i t does in fo rward or s ideward

f l igh t . Prope l le r power Increases , however, s ince t he prope l le r e ff ic iency i t se l f is reduced by t he reverse f low.

Figure 18 shows that , s ince the rudde r vanes are less e ffec t ive i n rearward f l igh t , more def lec t ion is required. Wnen fu l l def lec t ion was reached, fu r the r inc rease in rearward speed requ i red inc reased prope l le r pi tch .

FORWARD-SPEED CLIMB

Rate of cl imb versus a i r speed is shown in Fig u r e 19. The curves were not ex t rapo la ted beyond t he reg ion of measured data, and where curves were fa i red t h r o u g h l ess than f ive

points , they were drawn s u b s t a n t i a l l y para l le l to curves fa i red th rough f ive or more points .

The da ta presen ted at cons tan t a l t i tude (sea level s tanda rd ) , power, gross weight , and r o to r speed exhibi t the usual t rends with fo rward speed. hat is , ra te of cl imb inc reased to a maximum recorded value of 2300 feet per minu te f o r the f laps- up con f igu ra t ion at 80 knot s and then de- c reased with fu r the r Increase in airspeed. Power used in these t e s t s was 1050 horsepower, t he normal r a t ing of t he turbine. No t e s t s were made at mil i t a ry power (1250 horse - power) .

Of pr imary Impor tance is t he effect of a i rc ra f t a t t i tude at a g iven fo rward speed dur ing t he cl imb maneuver. y vary ing t he ro to r shaf t a t t i tude ( r e l a t ive to ver t ical ) f rom approx ima te ly -4 to +10 degrees ( f laps up) , the fuse lage ang le of a t tack var ied f rom approx ima te ly -5 to -15 degrees ( re la t ive to the f l igh t path) . The high ly nega t ive ang le of a t tack Is ind ica t ive o f the he l icop te r f l igh t t echn ique (nose down).

Since the pi lo t had no presen ta t ion of ang le of a t t ack , th i s i n f o r m a t io n became ava i lab le only a f t e r the da ta were reduced. As a resul t , some of t he angles of a t tack were

obta ined th rough only a l im i t e d speed range. here i s,

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however, an I n d i c a t i o n that t he he l icop te r t echn ique of c l imb is supe r io r at speeds be low 80 knots . n the other hand , t e s t s with f l aps fu l l down (HO degrees) and a fuse lage ang le o f a t t ack o f about +3 degrees re su l ted in t he highes t ra te o f c l imb obta ined (2380 feet pe r minu te at 67 kno t s . )

Rotor and wing l i f t in c l imb are shown in Figu re 20. As m e n t io n e d above , t he c l imbs were per fo rmed pr imar i ly by the h e l i c o p t e r t echnique . s a resu l t , the wing l i f t in c l imb is n ega t ive . Figure 20 shows tna t the effect of the f laps Is to reduce the nega t ive wing load , but that ang le o f a t t ack has l i t t l e e ffect (p robab ly because the ro to r

downwash i s t he dominan t i n f luence ) .

AUTOROTATION

Data were obta ined f rom steady a u t o r o t a t i o n descents f rom 19 to 101» kno t s wi th f l aps up and f rom H? to 102 kno t s with f laps down kO degrees. Land ing gear was extended in al l cases. Thro t t l e chops were conduc ted at speeds up to 139 kno t s , bu t the a i r speed s lowed so rap id ly f rom the higher speeds that i t was reduced be low 110 kno t s by the t ime t he needles sp l i t and s t eady a u to r o t a t i o n ha d occurred . i gure 21 shows the resu l t s of t he s t eady au to ro ta t ion t es t s in the form of ra te of descent versus a i r speed at cons t an t

gross weight and ro to r speed. Lowest r a t « ; of descent recorded was wi th f l ap s up, and was 131' 0 feet per m i n u t e at 58 knots . he ra te of descent i nc reased wi th speed unt i l i t » r a s approx imate ly 3200 feet per m i n u t e at lOH k n o f t s . he da ta do not y i e l d a clear m in im u m poin t because the photo panel malfunc t ioned d u r i n g several t r i a l s , but f rom p i lo t repor t s , the m in im u m appea rs to be between 50 and 60 knots . The f l aps-down ra te -o f -descen t curve is di sp laced f rom the f l aps-up curve, i nd ica t ing , as expected, tha t the a i rcraf t l i f t /d rag r a t io i s lower in the f l aps-down con f igu ra t i on .

The r a t e -o f -descen t da ta were al l ob ta ined at m i n i m u m prope l l e r pitch. An i n i t i a l objec t ive was to ob ta in addi - t i ona l d a t a wi th prope l le r p i t c h i nc reased 5 degrees. It was found, however, that the ro to r speed cou ld not be main ta ined a t or nea r the tes t va lue wi th a 5-degree increase in pi tch . The s t ab i l i zed ro to r speed at the higher prope l l e r p i t ch se t t ing cou ld no t be de te rmined because o f the engine governor l imi ta t ions descr ibed in the next paragraph .

Time hi s to r i es of sudden and complete loss of power could not be obta ined because of the ac t ion of the t u rb ine fuel con t ro l sys tem, which i s d iscussed under Technical Problems . However, t ime hi s to r i es wi th in the opera t ion l imi ta t ions of the t u rb ine g o v e r n o r were obta ined . Figure 23

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Figure 2 3 . Autorotatlonal Entry a t ^8.5 Knots.

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Rotor and w i n g l i f t are shown I n F igu r e 22 . i th flaps up , the rotor I s J J O percent unloaded at about 1 0 0 knots. Wi t h i n the speed range tested (45 to 1 0 0 knots) , f lap ex - tens ion further unloads the rotor an add i t iona l 1 0 percent of gross weight at a g ive n speed.

V E RT I C A L C L I M B

Ver t i ca l rate of c l imb versus power I s shown I n F igu r e 2 6 for tw o different referred gross weights . otor and prope l l e r power are d e r i v ed f rom Figures 25 and 12 , respec- t i v e ly. uro lne power i s d e r i v ed f rom th e combined f a i r ed curves of rotor and prope l l e r powerj corrected for mechanical losses as ex p l a i n ed under " D a ta Reduct ion Methods" . oints from the out-of-ground-effect hover curves (F lgurs 10) are used for the zero ra t e -o f -c l lmb points .

L E V E L F L I G H T

•

shows one at an entry speed of 4 8 .5 knots, and F igu r e 2k

shows one at a true ai rspeed of 139 knots. rom the behav io r of the ai rcraf t under the decreases I n power t y p i f i ed by these t ime h i s to r i es , the ind ica t ions were that no s ign i f i - cant differences w o u l d have occurred, except that the rotor speed would have i n i t i a l l y drooped to a lower va lue than that sus ta ined by the turbine governor. n the actual case, C as soon as th e turbine R P M dropped be low th e mi n i mu m governor se t t ing, the governor w o u l d cause the tu rb ine to supply enough power to mai n t a i n that R P M . n the 139-knot case (F igure 24) i t i s c lea r ly seen that reduct ion of prope l l e r • p i t ch not only arrested the R P M decay but caused i t to i n- crease momentar i ly s l i gh t ly above i ts i n i t i a l value. he w i n d m i l l i n g prope l l e r supp l i es power to the rotor and a lso creates addi t ional drag to decelerate the aircraft . n the 48-knot case (F igure 23) . the prope l l e r was already at es - s en t i a l l y min imum pitch, and th is effec t does not appear. A i r s p eed i s no t p lo t t ed i n Fi g u re 23, but i t was main ta ined at 48 to 5 0 knots.

Lev e l f l igh t power data are presented for var ious ranges * of fuselage at t i tude i n terms of prope l l e r and rotcr power coeff ic ients , as a func t ion of rotor advance rat io. esul ts ar e g i v en for two di f fe ren t test g ross -weigh t - to -dens l ty ra t ios: W /o « 6 2 0 0 pounds, and W /o « 6 7 0 0 pounds. ro- cedures set forth pr io r to actual f l igh t tes t ing s t ipu la ted that data be obtained at 20-knot In te rva ls , s tar t ing w i th 30 knots, for fuse lage at t i tude increments of +5 degrees to th e maximum t r lmmable at t i tude for each steady-state speed. Si n ce the p i lo t had no direc t presentat ion of angle of attack, hi s procedure was to f l y at var ious combinat ions of propel ler

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and co l l ec t ive pi tch . he re su l t ing f l aps-up da ta are presen ted as a f u n c t i o n of advance r a t io In l eve l f l igh t for ranges of fuse lage ang le of a t t ack of (1) -M to -6 degrees, Figures 28a, 28c, 29a, and 29c ; (2) -1 to +1 degree , Figures 28b, 28d, 29b, and 29d; (3) + ' » to +6 degrees . Figures 28b and 29b. Data f o r t e s t s wi th UO degree f laps down are given In Figures 28e and 2 9 e for an ang le of a t t ack of -0 .5 to +2 .5 degrees. The spec i f i c ranges of fu se l age angle of a t t ack were de te rmined (1) on the bas i s of the abundance of d a t a ava i l ab le with in an In te rva l and (2) because the midpo in t s of the In te rva l s are sepa ra ted by 5-degree Inc remen t s .

Sign i f i can t add i t iona l da ta were a lso produced In the ang le - of -a t t ack range of -1 .5 to -3 .5 degrees (Figures 28a, 28c, 29a , and 29c) .

The fa i red curves o f f l a p s - u p ro to r and prope l le r power coeff i c i en t as a f u n c t i o n of advance r a t io are compared fo r t he two gross weigh ts and the ang le -o f -a t t ack var ia t ions in the summary plo t s of Figures 2 8f and 29f . Two cons i s t en t fami l ies of curves r esu l t , exh ib i t ing the effects of both a t t i tude and weight on the component power coeff i c i en t s . These curves have been u t i l i zed to e s t a b l i sh t he dimens iona l , component power fo r ro to r and prope l le r based on s t andard atmosphere sea - leve l d e n s i t y and the refer red ro to r speed of

279 RPM shown on Figure 30 . Inc luded in Figure 30 are the f l aps-down componen t power curves ob ta ined f rom the d a t a of Figures 28c and 29e .

The equ iva len t f l a t -p la t e area of the bas i c a i rcraf t , as f lown, is ca lcu la ted to be 13 .53 square feet , as shown in Table XX in Appendix II . he ex t ra drag of t e s t - equ ip - ment i t ems and u n ln s t a l l e d fa i r i ngs o r ig ina l ly planned r a i sed the f l a t - p l a t e area for some f l i gh t s up to 1.2 square feet more. The net effec t of these i tems i s that the tes ts were conduc ted at an average d rag pena l ty of ^ O square feet more t han the c o n f i g u r a t i o n as designed . This ex t ra drag amounts to 50^ horsepower at 200 kno t s at sea level , and, combined wi th t he lack of ram recovery in the eng ine- in le t duct (equivalent t o 95 horsepower) , was the l i m i t i n g f ac to r on the l eve l f l igh t speed reached.

The h ighes t t rue a i r speed reached was 19 5 k n o t s in a 10-degree dive at an a l t i tude of 4 2 0 0 feet. ighes t true airspeeds reached in l eve l f l i gh t were 160 k n o t s a t sea leve l (F l igh t 411) , an d 167 k n o t s at 3670 feet (F l igh t 406) .

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of r o to r power t o t o t a l power was p l o t t e d on Figu re 33 s a func t ion o f ai rspeed for t he var ious ranges of fu se l age ang le of a t t ack wi th f laps neu t r a l and extended.

Figu re 3 | l shows the var ia t ion In r o t o r - l l f t - t o - g r o s s - welght ra t i o as a func t ion of t rue a i r speed with fu se l age * a t t i t ude cons idered as a parameter. he ro to r l l f t - to - weight r a t io s are somewhat higher for the heavier gros s w e i g h t , s i n c e the wing l i f t at a g iven ang le of a t t ack is cons t an t and the added weight must be ca r r i ed by the rotor. , The var ia t ion in ro to r l oad ing wi th ang le of a t t ack shown *

in the p lo t re f l ec t s t he i nc rease in wing l i f t wi th i nc rease In angle.

Rudder Po si t i o n In Level Fl igh t

Rudder p o s i t i o n i s » presen ted as a func t ion of t rue a i r- speed fo r va r ious fu se l age angles of attack. sepa ra t e curve Is p l o t t e d for each fu se l age ang le of at tack cons ide red , an d each p lo t shows the measured po in t s f rom which i t was cons t ruc ted (F igure 36).

These curves show that rudder def lec t ion changes in a di rec t ion tha t i s to be expec ted as a func t ion of a i r speed

and fu se l age ang le of at tack . he i nd iv idua l p lo t s fo r each fu se l age ang le of a t t ack show that the lef t rudder de- c reases as forward speed increases . n the l ower speed rang - es o f p t o approx imate ly 60 kno ts , wi th a fu se l age ang le o f a t t ack o f *\ to +6 deg rees , the r educ t ion i n lef t rudder ang le is large. his is due to the rap id reduc t ion in main ro to r power as fo rward speed i s i nc reased f rom hover t o approx imate ly 60 knots . In add i t i on , as a i r f l ow ove r the t a l l vanes i s i nc reas ing with i nc reas ing fo rward speed, less le f t rudder de f l ec t i on is requ i red to counterac t main ro to r torque. hi s effec t causes lef t rudder angle to con t inue to decrease wi th i nc reas ing fo rward speed even t hough main ro to r power s t a r t s to i nc rease as a i r speed increases beyond

the 60 knot range. he ra te of dec rease of lef t rudder ang le s ta r t s t o dec rease as ai rspeed increases beyond the 60 kno t range, due t o the effec t of i nc reas ing main r o t o r power an d * the rudder approach ing neut ra l .

Compar i son of the four curves for di fferen t angles o f a t tack shows that lef t rudder decreases as fu se l age ang le of a t t ack var ies f rom nega t ive to pos i t ive . his is as expected and re su l t s f rom the fact that , at a given speed, main r o t o r power decreases wi th i nc reas ing fu se l age ang le o f at tack .

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VIBRATION

Vlbration ecords ere btained t wiafeer f uselage locations hroughout he est perational ange. Suitable data or nalysis nd iscussion 1 ibration mplitude nd frequencies ere onfined o he ata btaineo 'c he ilot's station. (Actually, t as he opilot's tation, hich as structure dentical o he ilot's). Effort as oncentrated on his tem because f ts mportance n stablishing c- ceptable lying qualities.

Fourier nalyses of he easured ertical ccelerations

at he ilot's tation ere erformed n rder o etermine the mplitudes t iscrete otor anaonic requencies. This was one ecause ilot erception nd olerance evels epend upon requency. (M IL-H-8501A Reference 4) ives pper imits of cceptable ccelerations s unction f requency.) The extracted irst hrough he enth otor armonic omponents f the ertical cceleration re lotted n igures 38 hrough 47 s unction f orward peed. Included n ach igure s the cceleration rom he ata btained ith he riginal ontrol ystem} hat s, efore tiffening. This spect s iscussed n he Technical roblems'* ection, nd It s p- parent hat cceleration evels ere arkedly educed y tif-

fening, articularly t he hird otor armonic, hich s he largest cceleration omponent.

In IL-H-8301A t s tated hat ccelerations t irspeeds elow ruise peed re o e ess han .15g or .*e- quencies p o 2 ycles er econd through he ixth otor harmonic or he 6H-1A); t irspeeds bove ruise, ccelerations re o e ess han .20g or requencies p o 6 cycles er econd through he eventh cvor harmonic or he 16H-1A). At requencies igher han hesej, he pecification cites ibration imits n erms f isplacement, o hat he allowable cceleration s unction f requency. These l- lowable cceleration imits re hown n igures 38 hrough

47. Cruise peed s aken s peed or est ange, hich s 13 0 nots rue irspeed or he est ircraft.

Figures 8 hrough 7 how hat ibration mplitudes throughout he peed ange re ell elow he limits llowed by IL-H-8501A, ith he xception f he hird haraonic. The atter s ithin he pecification imit t peeds e- tween 0 nd 20 nots. It s lightly n xcess f his limit t peeds etween 30 nd 55 nots nd n he ransition ange etween 0 nd 5 nots. O ut oint btained n 195-knot, 10-degree ive h o w s hird - haraonic omponent f

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Vlbrational Frequencies —w^^w wTii , nn JHI .I I — ■ »in m ^i iii

The lbrational requencies f he ominant cceleration conponents ave een m easured at he otor ransmission main case Figure 8), t he propeller drive haft earing upport (Figure 9)» t he propeller haft (Figures 0, 1» nd 52), and at uselage tation *5 (Figure 3)« It s een hat he vlbratlonal requencies r e bout 36 ycles er econd or he transmission nd tall propeller drive haft bearing upport. The ail propeller haft otational peed s 5 evolutions per econd. The easured propeller haft requencies r e 40 cycles per econd n he ongitudinal, ertical, nd ateral directions, nd are hus loser o he ropeller otational speed. At uselage tation *5, he requencies re pproxi- m ately 2 ycles er econd. Three-per-rotor evolution s Ik ycles er econd.

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brations f he otor. There ppeared o e o ignificant

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FLYING A N D HANDLI NG QUALI TI ES A ND MANEUVERABILITY T h i s section contains t h e t e s t r e s u l t s a n d related a n -

a l y s i s , t h e o r y , and discussion on t h e 16H-1A flying and han- d l i n g qualities. t s major purpose i s t o compare t h e pertinent f l i g h t characteristics with t h e corresponding require- m e n t s a s given i n t h e Military Specification MIL-H-8501A ( R e f e r e n c e h). Throughout, t h e cited paragraphs refer t o t h o s e o f t h i s reference.

t I n addition, Appendix I I I contains t h e basic dynamic a n d « static stability and control t e s t data a n d analyses required

f o r t h i s s e c t i o n .

LONGI TUDI NAL CHARACTERI S TI CS

3 . 2 . 1 Reference M ) A sufficient margin of longitudinal control power i s r e -

quired s o t h a t 1 0 percent o f t h e maximum attainable pitching moment i n hover shall be available throughout t h e steady f l i g h t operating r a n g e .

From Figure 5 i t i s seen t h a t t h e hover trim position i s 1 . 9 inches forward, leaving 5 - 6 i n c h e s of forward displacement a v a i l a b l e . t 1 6 7 k n o t s , ( h i g h e s t speed reached i n l e v e l f l i g h t ) , t h e trim position i s 5 . 0 inches forward with a 2 . 5 i n c h margin remaining, a n d t h e pitching moment per i n c h of stick i s 7 7 percent of t h e hover v a l u e . h u s , t o m e e t t h e r e q u i r e m e n t , stick displacement available a t t h e high speed condition should b e

0 . 1 0 X 5 . 6 „ „ J . T 7 7— = 0 . 7 i n c h

H e n c e , sufficient margin i s available f o r t h i s center o f gravity position ( 9 . 6 i n c h e s forward). t t h e critical t e s t center o f gravity o f 3 . 1 i n c h e s forward, elevator trim a l o n e c a n b e made t o compensate t h e center of gravity shift f o r t h e s a m e stick p o s i t i o n . rom t h e data o f Figure 5 5 , which were obtained a t 1 3 0 k n o t s , i t i s estimated t h a t t h e

' elevator deflection p e r i n c h o f a f t center o f gravity movement i s 1 . 9 degrees per i n c h a t the high-speed condition. T h u s , 1 0 degrees o f d o w n elevator, o r 5 5 percent o f f u l l d e - f l e c t i o n , i s r e q u i r e d . xamination of a l l maneuvers p e r f o r m e d ,

I

ncluding autorotatlnn, quick s t o p s , rearward a n d sideward f l i g h t ( s e e Figures 1 0 2 , 6 0 , 5 6 , 5 7 , 8 ^ , 8 8 , and 9 ^ ) , shows t h a t c o n t r o l margin f o r a f t stick w a s a l w a y s more than a m p l e .

1 1 9

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i 3 . 2 . 2 Reference 1)

T h i s paragraph spec i f ie s a max i mu m longi tud ina l control motion of + 1 inch for s t i l l ai r prec is ion hove r i n ground effect . hi s test was performed i n F l i g h t 2 6 ^ , B l i p 10 i n w hic h the max i mu m mot ions recorded were less than + 1 . 0 inch ove r a recorded b l i p length o^ 17 seconds (see F igu r e 56) . Therefore , the r equ i rement i s sat is f ied .

3 .2 .3 . 3 . 2 . ^ t 3 . 2 . 6 , 3 . 2 . 7 and 3 .2 .8 L on g i tud ina l Contro l Forces.

(Reference * )

The ircraft, s ested, ad ully owered controls ith forces lways educed o ero (as erified y pilot's om- ments) ut ith o elf-centering. A n djustable ungee system dded o he ully owered ontrols ould atisfy he requireutent or positive elf-centering characteristics nd force gradient imitations.

3.2.3 (Reference

Tr a ns i t i ons f rom hover to an Ind ica ted airspeed of 70 knots i n 1 8 seconds and dece le ra t ions from 70 knots to hover i n 16 seconds were performed at constant a l t i tude withou t excess ive control motions and w i th greater var ia t ions of technique pe r mis s ib l e than i n conventional he l icop te r f l igh t . Records of two methods of acce le ra t ion were obtained. he least change i n at t i tude was by ma in ta in ing the sh ip l eve l l ong i tud ina l ly and inc reas ing the prope l le r pi tch to ac- celerate th e aircraf t in to forward speed. hi s w ould be the preferred method i n I P R opera t ions where the p i l o t w ou ld desire to I t e e p hi s instruments and at t i tude of the aircraf t to a mi n i mu m change. t ime his to ry of th is method at a forward center of grav i ty i s shown i n Figure 5 9« he qu ick stop from the second part of t h i s same b l i p i s shown i n F i g - ure 60 . Figures 6 l and 62 show cor responding maneuvers at an aft center of gravi ty.

T he convent iona l hel i cop te r mode of p i t c h ing down an d us ing the rotor thrust fo r accelera t ion i s i l lus t r a ted by th e t ime history i n Figure 63. pproxima te ly the same accelera- t ion was a c h ie ve d ( .26 g) , as measured from th e s lope of th e ve loc i ty - t ime curve. igure 6^ shows the quick stop w h ic h was performed immedia te ly afterward.

T he hor izonta l dis tance t r ave led during the t rans i t ions e i the r way was about 1 , 0 0 0 feet. he acce le ra t ing por t ion of the maneuver of Figure 5 9 was a ccompl ished almost en t i re ly w i t h th e propel ler. he at t i tude was never a l lowed to pi tch nose down more than 5 degrees. n th e other hand, as shown

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T I M E - S E C O N D S Figure 5 8. Ti m e H i s t o r y of Control Mot ions

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In Figure 6 3, the p ro p e l l e r pi tch was reduced as speed I n- creased, s ince the p ro p e l l e r was used on ly for ant i torque and yaw control .

A t h i rd method w o u l d be th e combina t i on of th e previous two, where some cy c l i c long i tud ina l s t ick motion forward i s g i v e n to i n c l i n e the rotor thrust , and s imul taneously prope l l e r pi tch i s increased to accelera te the ai rcraf t . he t ime of i n i t i a t io n of the maneuvers can be observed from both the propel le r p i t ch and the cy c l i c s t i ck motion.

3 .2 .9 Reference ** )

T he requi rement i s that angu lar accelera t ions be developed, i n the proper d i rec t i on , w i t h i n 0 .2 second after app l i ca t i on of long i tud ina l control for the speed range sp ec i f i ed i n paragraph 3*2 .1 . n examinat ion of the f o l l o w i n g control responses i nd i ca t es that th e 1 6 H - I A meets this requirement .

F l i g h t S peed Tes t B l i p F igu re

H o v e r 3 3 5 3 6 5 H o v e r 3 3 5 H 6 6 H o v e r 2 6 0 1 6 7 H o v e r 2 6 0 2 6 8 5 0 Knots 3 1 5 5 6 9 1 5 0 Knots ni ü 1 7 0 15 0 Knots 1 l H 2 7 1

3 .2 .1 0 Reference 4 )

T h i s paragraph states the requi rement for pos i t i ve control force and pos i t i on s t a t i c s t ab i l i t y as ev idenced by a pos i t i ve s lope of the l ong i t ud ina l s t i ck pos i t i on and force

versus speed curves.

moderate i n s t ab i l i t y i s a l l owed at forward speeds of 15 to 5 0 knots an d at rearward speeds of 10 to 3 0 knots. i th a power-boost system, i n conjunct ion w i t h a r t i f i c i a l feel, the force s t ab i l i t y w i l l be pos i t i ve i f th e p o s i t io n s t ab i l i t y i s p o s i t i v e .

T he 1 6 H - 1 A possesses p o s i t i v e c according to th e requirement . hi s f l igh t s t ick pos i t i on versus speed th e s lope i s p o s i t i v e from about 70 speed, w i t h a s t ick reversal of O.k

50 knots. aragraph 3 .2 .1 0 a l l o w s 0 .5 inch i n this speed range. he obta ined us ing the s t i f fened contro Problems") and are s ig n i f i can t ly di

ontrol pos i t i on s t ab i l i t y i s seen i n th e l eve l

plot of Figure 5 ^, where knots to th e m ax im u m inch between hover and

a m ax im u m reversal of data for this curve were 1 system (see"Technicai fferent from the i n i t i a l

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resul ts obta ined , which showed a curve of t he same genera l shape but with s l cpea 70 t o 30 percent g rea ter. In genera l , abso lu te values of p o s i t i o n are reduced as a resul t o f t he control modi f i ca t ion .

The s p e c i f i e d s p e e d range f or p o s i t i v e s t a t i c s tab i l i ty in au to ro ta t ion is f rom 60 to 100 percent of the m a x im u m au to ro ta t ive speed. hi s r ange Is covered in the da ta of Figu re 102 where a p o s i t i v e s lope for both f l aps-up and f l aps-down c o n d i t i o n s i s shown. he c l imb c o n d i t i o n fo r which t he s l o p e must be pos i t ive is t he speed fo r m a x im u m ra te of c l imb +15 knots . i nce the normal ra ted power c l imb curve shown in Figure 72 has a p o s i t i v e s l o p e at speeds above 65 knot s , and s ince speed fo r m a x im u m ra te of c l imb i s 75 t o 80 kno ts» th i s r equ i r ement is sa t i s f i ed .

The curves of Figures 5 * * and 72 are based on da ta f rom many di ff e ren t f l i gh t s . In orde r to i nves t iga te the s t a t i c l o n g i t u d i n a l s tab i l i ty more c l o se ly, f l i gh t s were per fo rmed i n which the a i r c r a f t was t r im m e d t o a p a r t i c u l a r speed in level f l i gh t , and a b l ip was t aken . hen, h o ld in g a l l o the r con t ro l s f ixed , t he l o n g i t u d i n a l s t i c k was success ive ly t r immed to new speeds (not necessa r i ly l eve l f l i gh t ) +10 knot s and +20 knot s f rom t he i n i t i a l t r immed l eve l - f l igh t

Speed , and b l i p s were t aken at each t r im po in t .

he same

procedure was used in c l imbs , where t he i n i t i a l t r im po in t was in a c l imb at normal r a t ed power ra ther than l eve l f l ight . This procedure i s in s t r i c t accordance wi th Reference 4 .

Resul ts of these t e s t s are shown in Figure 73 fo r i n i t i a l t r im speeds o f 60 kno t s in level f l igh t nd 50 knot s in climb. Reference t o Figures 74 and 72 shows tha t at these speeds the 16H-1A has leas t s t a t i c s t a b i l i t y in l eve l f l ight and c l imb, respect ively. Figure 7 ' » , in which the curves of Figure 73 are s u p e r im p o s e d on t hose of Figures 5^ and 72 , shows somewhat g r e a t e r s t a t i c s t a b i l i t y when t he maneuvers are per fo rmed in the presc r ibed manner, even wi th a r ß o r e af t

center o f grav i ty.

3 . 2 . 1 0 . 1 Reference 4 )

This parag raph re fers to the opera t ing cond i t ions to be cons ide red i n 3-2 .10 . Al l s t eady f l i gh t s fo r more than sho r t t ime i n t e rva l s are to be considered, as wel l as the most cr i t i ca l cen te r-o f -g rav i ty pos i t ion . he da t a refer red to in t he preced ing parag raph cover al l cond i t ions spec i f ied in MI L- H - 8 5 0 1 A except par t i a l power descen t s , fo r which da ta were not requi red . Cons ide r ing the overa l l s t a b i l i t y charac te r i s t i c s o f the 16H-1A there is no i n d i c a t i o n o f in- s t a b i l i t y fo r th i s o p e r a t i n g cond i t i on . es t s at centers

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Pos i t ion versus Speed, C.G. 8.6.

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•

of grav i ty between 9 .6 Inches and 3 .1 inches forward of the shaft do no t ind ica te any s ig n i f i can t a l t e ra t ion of the s t ab i l i t y and control der iva t ives .

^ . 2 . 1 0 . 2 Reference 4 )

T he long i tud ina l control v a r i a t io n s for t r im i n c l imb , l evel f l igh t , or part ia l power descent at a constant speed w i t h i n th e f l i gh t enve lope are sp ec i f i ed to be not more than + 3 inches. T he speeds mentioned as need ing th e most sp ec i f i c i n v es t ig a t io n ar e Vraax and speeds between zero and one-half th e speed for m i n i m u m power (5 0 knots for the 16H-1A) .

Refer r ing to th e curves of Fig u re s 51* and 72 for l eve l f l igh t and c l imb , respec t ive ly, i t i s seen that at zero forward speed, the apparent s t ick pos i t ion change i s 3-^ inches, and i t becomes less than 3 inches at speeds between 3 0 knots and 80 knots. t 10 0 knots, the ind ica ted change between c l imb and l eve l f l i gh t i s 3.5 inches. I t should be rea l ized that the c l imb data were obta ined for a center-of- grav i ty p o s i t io n of 3.1 inches forward, and th e l eve l f l i g h t resul ts are for a center-of -gravi ty pos i t ion of 9 .6 inches forward. he cen ter-o f -g rav i ty difference wo u ld tend to ind ica te sm a l l e r changes when cons ider ing the same center- of -grav i ty pos i t ion .

Although part ia l power descents were made on many occasions, data thereon were no t required under th e contract, and were no t obtained. owever, compar ison of data for full autorota t ive descents (from Fig u re 1 0 2 ) w i t h c l imb data (from Figure 72) shows a maximum t r im change of 5 inches at 70 knots between zero power descent and normal rated power cl imb. gain , part of th i s t r im change can be at t r ibuted to a cen ter-o f -g rav i ty diffe rence of 6 inches.

3 .2 .11 (Reference 4 )

T h i s sp ec i f i ca t io n f i x e s the long i tud ina l d y n am ic sta- b i l i t y requirements . I t i s stated that sat isfactory dynamic s t ab i l i t y charac te r i s t ics shal l be exhib i ted . p e c i f i c a l l y, th e short-per iod and phugoid mode m i n i m u m damping requirements are g i v e n as a function of the o s c i l l a t i o n per iod . These s t ab i l i t y l i m i t s i n terms of th e t ime to damp to h a l f ampl i tude or th e t ime to double ampl i tude , f o l l o w i n g lo n g i - tudinal disturbances, are presented i n Fig u re 76; the corresponding spec i f i ca t ion paragraph i s noted. hey appear as shaded l i m i t l i nes denoting the m i n i m u m damping, and they have been made functions of inverse t ime i n order to obtain a continuous boundary over th e en t i re per iod range.

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In cmU' of he aplace *r:.i: ." »* | .? .omP'-9

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An .325 .ivgs" 1 -5 2 * ir,, ... C^3)

Equating he econd erivative f n o >vro s rdc.?- o determine he nflection-point ime ields

^r^ (1.52) 2 (3.18) 2 1 in 3.l8t .57)

(2 .52 .18) O S (3.i8t ,57)

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' -., ; '-. - omal * $ itr st ion esponse c ea gmjz./e >;;«•:, -: • , > üft r .ê :,art f he aneuver, nd lnce

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I h« **-..- .• ,.;.. . .u- -. ^ >^ftii lso ecorae or.eav« ow n I Jtu'ft • - . -.-t'- ;w fivi -f he Ba,neuv*r. ««-re ll | '.âî .^.;-v-^ -^-\;CX'..£. i-v-fc %re e e onsidered. Because

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v f :C lavo.';- s «ai,y-A£?c or his ase.

l-m; itch-rate esponse quation o ontrol o^Jnr; 3 '

iven by

where K M 6

1 a , M a Y L i iV ^ mV ,

and the r emain ing parameters are as p rev io u s ly defined.

or a 1 - ln ch af t step control disp lacement , and wi th the 5 0 - knot d e r iv a t iv e s (as g i v e n i n Appendix I I I ) appl ied , th i s equat ion i n terms of th e Lap lace v a r i ab le and i n degree-per- second units i s

*( s) 3[ Jt ffrrM .59J2] >

and the t ime so lu t ion , or Inverse transform, i s then

6 • ^-65 - 5.-«0e- 0 ' 9 1 t s in (1 .3 1 t + 2 . 1 0 ) (48)

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Equa t ing the s tconü der iva t ive of e to zero in o r d e r to es tab l i sh t he in f l ec t ion -po in t t ime yie lds

f-i » 1(0 ,91 ) 2 - {1.31) 2 ] s in (1 . 31 t + 2 .10 ) -

( 2 X 0 . 9 1 X 1.31) cos ( 1 .3 1 t + 2 .10 ) » 0 {^9)

or

an (1 .31t +2.10) = 2 .69

and

. 2 6 sec (50)

I

hus , under the most cr i t ica l condi t ion , the i n f l ec t i on -

| oin t t ime exceeds t he s p e c i f i e d maximum, bu t only by 13 I ercent . ince the response is a damped second-o rde r o s c i l - | at lon, i t must r * f ~ i a i n concave down f rom the i n f l e c t i o n po in t I o t he maximum p i t ch rate.

3 .2 .11 .2 (Reference 1 )

This paragraph concerns t r i e c a p a b i l i t y of Inherent

l imi t ing o f

acce le ra t ion

re sponse

due to shor t

d u r a t i o n

d i s -

tu rbances . p e r a t i n g speeds are not spec i f led , but t he h igh- speed (150-kno t ) case is f i r s t analyzed . t is s t a t e d tha t when t he s t i ck mo t i o n is a 0 .5 - s e c o n d af t square pu l se ( to s imula te a d is tu rbance ) the change i n normal a c c e l e r a t i o n magn i tude sha l l not exceed 0 .25 g a f te r the c o n t r o l has been re tu rned to t r im. he cont ro l input ma g n i t u d e is to be the leas t of: 1 inch , the s tep s ize requ i red to produce a 0 .2 radian per second p i t ch ra te with in 2 seconds , or the s tep s ize requ i red to deve lop a normal a c c e l e r a t i o n ( to ta l ) o f 1.5 g with in 3 seconds . Co n s i d e r i n g s t e a d y - s t a t e va lues , i t is found f rom the re sponse equa t ions g iven in t he preced- ing paragraphs tha t t he s t i ck def lec t ion for the p i t c h - r a t e

cr i te r ion is

6 1' -^MJ 7'3

5 1 )

= - 2 .46 inches

For t he normal acce le ra t ion cr i te r ion , the s t i c k d e f l e c t i o n is

62 = - Qjj - 11

52)

^ 5 7 3 2 5

= - 1. inches

155

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w i th re sponse t imes less than the maximum t imes ci ted. hus, a 1- lnch af t square pu lse Is adop ted , and for the purposes of analys i s , th is re sponse can be c lose ly app rox ima ted by an Impulse Inpu t o f -0 .5 Inch- second .

Since on ly acce le ra t ions a f t e r con t ro l is r e tu rned to t r im need be cons ide red , the direct con t ro l con t r ibu t ion to a c c e l e r a t i o n Is omi t ted and only that due to ang le o f a t tack is computed. I t can be seen f rom t he r esponse equa- t i o n of paragraph 3 .2 .11 .1 (a) that , for a -0 .5 inch -second i m p u l se i npu t , t he r e sponse equa t ion in the Laplace var iab le Is

«

(An)a 3 ) -0 .5 X 0 . 4 0 5 ( s -2 .28)

[s 2 + 2(Q.H30) (3 .52 ) s + (3 .52) 2 ](53)

and the t ime so lu t ion i s then

(An)a - -0 .315 e- ^sin (3 . l8 t + 5 .60) (5 M

E x amina t ion o f t h i s re sponse Ind ica tes that a nega t ive peak of -0 .120 g is reached at 0 .568 second^ ucceed ing peaks are smal le r in magn i tude as the r esponse is damped.

At a 50-kno t fo rward speed, the acce le ra t ion r esponse due to a is l a rge r than at 150 knot s , whereas that produced by direct con t ro l con t r ibu t ion i s sma l le r ( -0 .07 g/ inch compared t o - 0 . 2 5 g/ lnch) , so that th i s peak acce le ra t ion change s h o u ld be Inc reased accord ing ly. I f a cor rec t ion propor- t i ona l to the r a t io of s teady s t a t e (An)a f rom a s tep input at t he two speeds i s app l i ed .

peak (An) 0 .13 t 4 * ( -0 .12) (55)

- -0 .215 g

This is less than the max imum a l lowab le of 0.25 g and hence t he s p e c i f i c a t i o n requ i remen t is sa t i s f i ed .

3.2 .12 (Reference 4 )

The requ i remen t is that , fo r the maneuver of parag raph 3.2.11.1 , the normal a c c e l e r a t i o n always increases with t ime unt i l the maximum i s at ta ined. As the bas ic acce le ra t ion re sponse es tab l i shed in the preced ing analyses is damped second-order re sponse and akes place only in t he shor t -

1 5 6

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per iod mode, i t I s apparent that acce le ra t ion reversa l as descr ibed i n t h i s paragraph does not occur.

3 .2 .1 3 Reference H)

Long i tud ina l control power i n h o v e r at m a x i m u m o v e r lo ad weight or rated power shal l be such that a 1- inch step input from t r im w i l l produce, w i t h i n 1 second, a m i n i m u m p i t ch change of

6 s V" ^ l o o o ' degrees

T he test weight of 6 0 0 0 pounds i s app l i ed . ence,

. 5

(56 )

{J 0 0 0 + 1 0 0 0

=2.35 degrees

(57 )

R efe r r i ng to the hover pi tch-ang le response equat ion presented i n Appendix I I I , t h e p i t ch angle response to a 1- inch af t s tep i s g i ven by

i(s) 7 .8 0 ( s • > • 0. 0 0 6^ )

8 ( s 0 . 6 2 2 ) [ s2 + 2 ( -0 .3 0 5 ) (0 . i < 0 3 ) s + (0.i»03) 2 3 (58)

and he orresponding nverse ransform or ime olution s

e .50 l.r'»- 0 ' 6 2 t +2i».ile0 -1 2 3 t s l n(0 . 3B5t .80)

F rom this equat ion, at t « 1 second, « 3 .36 degrees. Therefore the requirement i s sa t i s f i ed T or th e test weight . At h ig h e r gross weights, th e requi rement w o u l d be sa t i s f i ed by an increas ing margin , s ince control power i n h o v e r i s es sen t i a l l y proport ional to weight , w h i l e th e required response becomes less as weight increases .

T he paragraph a l so s t ipu la tes that when m ax im u m a v a i l a b l e displacement from t r im of the l ong i t ud ina l control i s app l i ed , th e pi tch angle sha l l be at leas t 4 t imes as great as the pi tch change fur a 1 - l r . c h displacement . T he control p o s i t io n curve for tn e 1 6 H - I A reveals that t r im i s at 1 .9 inches forward i n trver. i nce 5 .6 inches of t ravel remains, and a l i nea r ana lys i s approacu i ^ appl icable , th i s requi rement i s also sat is f ied .

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w i th af t cycl ic con t ro l were required to s top the ai rcraf t .

2 . Figure e igh t s and box pa t t e rns were acccmpl l shed at g r o u n d speeds o f up t o 10 knot s wi thou t the use of brakes. Above 10 kno t s In the above wind cond i t ion , brakes and/or increased co l lec t ive pi t ch were required.

3 . Main r o to r blade c learances were ample at al l t imes with 100 percent ro to r RPM and m l n i r a u r a co l lec t ive p i tch . Rotor engagements were a c c o m p l i s h e d withou t d i ff i cu l ty in the above winds .

' J . The a i rc ra f t cou ld be s topped at any t ime while p e r f o r m in g t hese maneuvers by the a p p l i c a t i o n o f brakes or cycl ic con t ro l and the add i t ion of co l lec t ive pi tch .

The con t ro l p o s i t i o n s ( co l l ec t ive pi tch , t a i l p r o p e l l e r pi tch , t a i l vane se t t ings an d r o to r cyc l i c p i t ch ) were moni tored by t he p i lo t and cop i lo t -eng inee r dur ing most of these t es t s . s c i l l o g r a p h records were ob ta ined for two f igu re -e igh t maneuvers i n 10-kno t winds . n g i n e RPM was

he ld cons tan t by t he eng ine governor.

A camera was used to reco rd the clearance between the ro to r t ip pa th p lane and t he t a i l duct , which never was less than 41 inches .

The f o l l o w in g techniques were developed:

1 . When t ax i ing upwind or crosswind , the t a l l p r o p e l l e r pi t ch was used fo r acce le ra t ion and dece le ra t ion a id the rudders were used f or d i rec t iona l cont ro l . The cyc l i c s t i c k was he ld neutra l fo r acce le ra t ion , back f o r dece le ra t ion , and i n to the wind at all

t imes du r ing c rosswind t ax i ing .

2 . When t ax i ing downwind , co l lec t ive pi t ch was increased and cycl ic con t ro l was used fo r acce le ra t ion and dece le ra t ion .

All o f the above t ax i ing was repeated at 93 percent main- ro to r RPM (5600 eng ine RPM) withou t di ff i cu l ty. Dece le ra t ion was more eas i ly per fo rmed under these power cond i t ions . Figures 82 and 83 i l lu s t ra te th i s capab i l i ty i n the recorded da ta for f igu re -e igh t maneuvers in 10-kno t winds.

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3.3-2 Reference t )

Figure SM presents the required control pos i t ions for s ideward f l igh t i n both di rect ions up to speeds of 30 knots , th e contract requirement .

Marg in of Control Power

30-Knot Lef t S ideward Fl i g h t

A n examina t i on of Fi g u re 8^ reveals that the m a x i m u f f i control pos i t ions are Inches af t s t ick , 2 inches l a tera l s t ick» and 1 inch r ight pedal . hus, control disp lacement margins for longi tudinal s t ick , lateral s t ick v and pedal respect ively are ^7 percent, 69 percent, and 69 percent of t ravel from neutral .

30-Knot Righ t Sid eward F l i g h t

Fi g u re S * » also ind icates maxiamms for t h i s f l igh t condi t ion of 3.5 inches forward s t ick , 2 Inches r ight s t ick , and 2 .5 inches le f t pedal . he resu l t ing control marg ins are then 5 3 percent l ong i t ud ina l s t ick , 69 percent l a tera l s t ick , and 23 percent pedal .

C ons ide ra t i on should a l so be g i v e n to the data contained i n Figures 85 through 96 i n connect ion w i t h s ideward f l i g h t control margins. I t i s ev iden t i n these t ime h i s to r i e s that r o l l , pi tch , and y aw maneuvers du r ing s ideward f l i g h t i n each d i rec t i on were s u c c e s s f u l l y performed.

3.3 .3 Reference ^ )

Hover over a g i v e n spot should be p o ss ib l e w i t h no more than + 1 inch of la teral and d i rec t i ona l control motions. Hover data for Fl i g h t 2 6 4 , B l i p 10 (F igure 5 8 ) , Ind i ca t e that lateral s t i ck mot ions were less than + 0 . 7 5 i nch and pedal mot ions were less than + 0 . 6 0 inch , for the 17-second dura t ion of th e bl ip .

3 .3 .^ Reference 4 )

I n al l steady opera t ing condi t ions , su ff i c i en t l a tera l control margin amount ing to at least 10 percent of the m a x i - mum hover control moment shal l be av a i l ab l e .

T he hover la teral t r im pos i t i on i s 1 inch le f t (F igure 97), l eav ing a m ax im u m lef t control d i sp lacement f r o n i t r im of 5 .3 inches. he control power, based on the ra t io of

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con t ro l moment per inch to iner t ia , I s 5 1 . 0 degrees/second 2

p er inch I n hover (Appendix I I I , page 3?^) . hl c g i v es a requ i red marg in of at least

0 . 1 0 X 5 .3 X 5 1 . 0 - 2 7 degrees per second 2

The c r i t i ca l opera t ing cond i t ion i s at h igh speed (150 knots) where the t r i m pos i t ion i s 2 .3 inches le f t (Figures 96 and 99). t th is speed, lateral control power i s 35 .8 degrees per second per second per inch (Appendix I I I , page 367) . Thus the m ln lm u x av a i l ab l e control i s

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3 .3»5 Reference 1 )

Direc t iona l control power shou ld be such that i n the hover cond i t ion a 1 - lnch step pedal input w i l l produce a yaw ang le di sp lacement at the end of 1 second of at least

1 1 0

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»5.71 degrees

In hover, t he yaw response to a 1- inch pedal step dis- p l c c c » . < ; n t is computed for the 1 6 H - 1 A f rom the di rect ional

r esponse equa t ion g i v e n i n Append ix I I I , page 3^5, equa t ion (81). his so lu t ion has been v e r i f i ed by dynamic response correlat ions as descr ibed i n th e append . n the Lap lace va r i a b l e , the equat ion for a unit step i s

* K ( s + |)

(63)

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Wi t h these parameters , th e response so lu t i on for a 1- inch lef t-pedal step displacement i s

i = - 5 7 0 ( e - * - 1 5 9 1 + 0 . 1 5 9 t - l ) 61 )

Th i s response ind icates a yaw motion at the end of 1 second of - 6 . 8 3 degrees, which i s approx imate ly 2 0 percent better han th e m i n i m u m requirement . F or a r igh t -peda l displacement, the ya« angle at the end of 1 second i s 1 5 0 percent oetter than he sp ec i f i ed value , s ince i n t h i s case

K - 3 0 . 0 degrees per second per second, as g i v e n i n Append ix I I I .

When m a x i m u m a v a i l a b l e pedal motion from tz la i s appl ied , the yaw angle at th e end of 1 second should be 3 t imes the 1- inch requirement . T he 1 6 H - 1 A was tested i n accordance wi th contract requirements, wi th the "most cr i t ica l t a i l - propel le r-p i tch /vane combina t ion" , so that only 1- inch l e f t - pedal t ravel from t r im remains, wh ich i s not the normal p i l o t technique. i th th e combina t ion of p ro p e l l e r pi tch and vane d e f l ec t io n used i n normal technique, th is r equ i rement can be met.

3-3.6 (Reference 1 )

I t should be poss ib l e to execute complete turns i n h o v e r i n a 35-knot wind. I n a steady h o v e r at the most c r i t i c a l azimuth angle r e l a t i ve to the wind , fu l l pedal d i sp lacement i n th e c r i t i ca l d i rec t ion should produce a y aw disp lacement (for the 1 6 H - 1 A ) of at least 5 .7 ^ degrees at the end of 1 second.

Complete h o v e r in g turn data i n both direc t ions ar e shown i n F igu res 10 0 and 10 1 for a 15-knot w i n d ve loc i t y. T he s t ick posi t ion shown I s wel l w i t h i n l i m i t s ; however, i n th e le f t turn, ful l le f t pedal was required. However, aga in i t i s noted that th e maneuver was performed wi th th e "most cr i t i ca l t a l l - propel le r-p i tch /vane combinat ion" , wh ich i s not the normal conf igura t ion of pi tch an d vane def lec t ion . Using a more favorable combinat ion , th e requirement can eas i l y be met i n a 15-knot wind , and probably i n the required 35-knot wind as w el l , al though such winds di d not occur dur ing the test period.

3 .3 .7 Reference 1 )

T h i s paragraph spec i f ies a max imum direc t iona l control power value i n h o v e r as that wh ich produces a 50-degree y aw disp lacement at the end of 1 second fo l lo win g a sudden 1 - in ch pedal displacement .

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A t t h e lightest normal service loading ( a s included i n t h e specification), t h e estimated ratio of control m o m e n t t o inertia i s 3 0 . 0 degrees per second per second per i n c h f o r right-pedal i n p u t s . ased o n t h e results contained i n t h e paragraph 3 . 3 . 5 analysis, t h i s value will produce a yaw change of 1 6 degrees a t t h e e n d of 1 second f o r a 1 - i n c h pedal step i n p u t . h u s , maximum directional control i s well within t h e l i m i t .

3.3.8 Reference Ü )

Although coordinated turn t e s t s were made i n autorotation, data thereon were not required under t h e contract a n d w e r e n o t obtained. owever, based o n t h e near-neutral stick posi- t i o n s and v a n e deflections shown i n t h e control position v e r s u s autorotative speed c u r v e s o f F i g u r e 1 0 2 , t h i s require- m e n t posed no difficulty.

3 . 3 . 9 Reference k)

This paragraph c o v e r s t h e general requirement f o r posi- t i v e static directional stability and positive effective dihedral, a s indicated b y a positive s l o p e o f t h e pedal- displacement and lateral-stick position versus sideslip c u r v e s . he 16H-1Ä static directional characteristics a r e analyzed i n Appendix 1 1 1 with t h e t e s t data a s shown i n Figure 1 0 3 . e r e , positive static stability i s indicated from a minimum forward speed of 5 0 k n o t s ( t h e l o w e r requirement l i m i t ) t o high s p e e d . t l o w speed, a sideslip a n g l e r a n g e of + 2 5 degrees w a s c o v e r e d .

*

Essentially knots b y a c o n s t degree sideslip a n g l e range d i d

speed dihedral e near zero ( g e o m e effects); howeve i n g . ence, a t 16H-1A possesses

z e r o dihedral e f f e c t w a s indicated a t 5 0 a n t lateral-stick position over t h e + 2 5 - r a n g e . t 1 0 0 k n o t s , the limited sideslip n o t permit conclusive evaluation o f t h e high

f f e c t . he 16H-1A wing contribution i s t r i e dihedral p l u s t h e wing-fuselage Juncture r , t h e rotor contribution i s a l w a y s stablliz- a n d a b o v e 1 0 0 k n o t s , i t i s estimated t h a t t h e

a positive l e v e l o f effective d i h e d r a l .

3 . 3 . 1 0 , 3 . 3 . 1 1 . 3 . 3. 1 2. 3 . 3 . 1 3 . 3 . 3 . 11* R e f e r e n c e 1 )

Lateral and Directional Control Forces

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j * 3 .15 (Reference k) 3

T h i s parafrraph speci f ies an upper l i m i t on response to lateral control def lec t ion at al l speeds. he m a x i m u m rate of ro l l per Inch of lateral s t i ck disp lacement I s no t to

ü s l n s only lateral cyc l i c control as p ro v id ed i n the research ai rcraf t , i t was necessary to have a h ig h degree of control power i n hover i n order to have adequate control at V j r ^ j f . Control modif ica t ion i n v o l v i n g reduced l a t e ra l - s t i ck -eyc l i c gear ing i n conjunction wi th a i l e rons (d i fferent ial f laperon def lec t ion) operated by th e s t ick wo u ld y i e l d m ax im u m rol l ^ates of less than 20 degrees per second and yet o* " suff ic ien t magni tude at ai l opera t ing speeds.

3.3 .16 (Reference h)

T h i s paragraph l i m i t s the delay i n th e deve lopment of angular ve loc i t y to la teral or direc t iona l control to 0 .2 second.

I i

An examina t i on of th e la teral and direc t iona l control responses conta ined i n Appendix I I I (Figures 1 9 ^ , 195 , and 1 98 through 203) ind icates t^at th i s requi rement i s sa t i s f ied ,

3 .3 .17 (Reference k)

T he hel icop ter shal l not ex h i b i t lateral t r im changes i n excess of 2 inches w i t h charges i n power and/or co l l ec t iv e pitch.

A n examina t i on of control pos i t ions i n steady ver t ica l c l imb at various co l l ec t i ve and power se t t ings (F igure 1 0 ^ ) shows that the lateral t r im pos i t i on i s essen t i a l ly constant over the test range. n Figure 105 lateral s t i ck p o s i t io n

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I s p lo t ted against bpeed for l eve l f l igh t , c l i mb and autorotation. T h i s f igure shows that throughout the range tested la tera l t r im does not vary more than the a l lowed 2 inches .

3.3 .18 Reference 1 )

Latera l control power i n hover sha l l be such that a r ap id 1- inch la tera l control step displacement shal l produce an angular change at the end of 0 .5 second of at least

21

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T he 1 6 H - 1 A response i s determined from the tes t -ver i - f ied character is t ics g i v e n i n Append ix I I I (page 353) . he uni t step response i s g i v e n by

♦ ( s ) = K

^1 ( 6 6 )

where, based on the 50 - kn o t dynamic response corre la t ion,

K = 53 .8 deg/sec 2 per inch

5 ? » 1 .38 sec 1

Wi t h these va lues ,

th e response so lu t ion i s 0 = 28 .2 (e- 1 ' 3 8 t + 1 .38 t - l )

Set t ing t = 0 .5 second i n th i s express ion yie lds a ro l l angle of ♦ « 5 . ^ 0 degrees. Thus the 16 H - 1A more than f u l f i l l s th e spec i f i ca t ion requirement by a factor of 3-8. eferr ing back to paragraph 3 . 3 . 1 5 , I t was shown that the rol l s en s i t i v i t y i n hover exceeded the maximum permit ted i n that paragraph by a factor of 1.9. Hence, cutt ing th e l a t e ra l - s t i ck swash-plate gear r a t io i n h a l f would resul t i n meet ing that requirement w h i l e s t i l l s a t i s fy i n g th i s one.

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of he ircraft s pproximately evel t ina _ the yclic stick may e held n he ame ongitudinal position, lthough some ft longitudinal yclic will id n nducing he irspeed nd maintaining or ncreasing he otor P M . The collective itch s of econdary mportance n ntry nto autorotation rom V^x because t s already t ow setting. Therefore, t ay e educed after he nitial deceleration or after he irspeed s educed o 00 o 120 nots.

For east rate of escent, a 60-knot glide ith minimum

tail propeller pitch nd minimum collective pitch hould be stablishsd.

A tep-by-step rocedure follows:

1. educe ail propeller pitch.

2. ecelerate, olding light ft ongitudinal cyclic stick.

3 . educe ollective itch o minimum at 10 0 to 25 nots.

M . Establish Q-knot glide.

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STRESSES AND LOADS

OBJECTIVES

The pr imary reason fo r measuremen t of loads and s t resses f

was t o assure s a fe ty of f l igh t In the conduct of the t es t program. Fo r th is reason , the ob jec t ive was to ga ther data dur ing a l l reg imes of f l igh t n ro to r and p rope l l e r bending

moments , both f l apwlse and chcrdwlse , shaf t to r s ion , ro to r • shaf t ben d ing and t ens ion , and ro to r and p rope l l e r blade con- • t ro l moments .

DESCRIP TION OF TESTS

On e ro to r blade was ins t rumented to read f l apwlse bending at s ta t ions 46. 5 9 . 5 , 79.2 , 105.6 , 124, 132, 158.4 , 184.8 , 211*2, end 237.6 , he choice o f these s ta t ions for measurement of f l apwlse bend ing was based o r . seve ra l factors as fo l lows:

1 . o have gages spaced c lose enough toge ther (every

10} of blade r ad ius) to give a reasonab ly clear p ic tu re of blade f l apwlse moments and s t re sses a long t he en t i re l eng th of the blade.

2 . o de te rmine the f l apwlse moments and s t resses at loca t ions on t he blade s p a r where there are t r ans i - t ions f r o m one size and/or shape cross sec t ion to another. Ins tances of th is are s t a t ion 46, which is a round sec t ion ; s ta t ion 5 9 . 5 » which is in a t rans i t ion area Jus t inboard of the f i r s t t r a i l i ng edge box; and s ta t ions 124 and 132, which are at each end a t r an s i t i on sec t ion on t he D spar.

3 . o use 1 . at lons for gages prev ious ly used on t e s t s of the H-21 metal b lade fo r purposes of compar ison.

St ra in gages were i n s t a l l ed on the ro to r shaf t to measure ro to r l i f e ( t ens ion) , bend ing moment in two perpendicu la r p lanes , and t o r s ion .

On the t a i l prope l le r, s t ra in gages were Ins ta l led to . measure hub bending; f l apwlse bend ing s t resses on one blade ac 40, 50, 70, and 90 percent blade span; and prope l le r p i t c h l ink load.

1 9 4

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Referring o Figures 07 hrough 26, he lots f teady and lternating lapwlse om ents nd tresses ersus lrsoeed show, n eneral, hat he teady tresses n he otor blade inboard f tation 32 ecrease lightly nd utborad f tation 32 ncrease lightly, with n ncrease n orward peed. The lternating lapwlse om ents nd tresses nc.ease radually with n ncrease n orward peed.

Variation f otor-blade lapwlse om ents nd tresses

along he lade pan or ypical lights t irspeeds f 0, 120, 50, nd 95 nots how hat nboard, he teady lapwise om ents re hose which ut ension n he ottom urface f he otor blade, nd utboard, he teady lapwlse moments re hose which ut ension n he op urface f the otor blade. The oint f nflection, or ero teady moment , oves nboard s he peed f he ircraft ncreases (see Figures 27 nd 28)* The otor-blade lapwlse alternating bending moments eak t 0 ercent adius t all irspeeds lotted, rom 0 nots o 93 nots.

Chordwise om ents nd tresses lso how definite at- tern. T he teady hordwise om ents nd tresses ecrease

with n ncrease n irspeed t tation 6 nd t tation 131.5» hereas he lternating hordwise om ents nd tresses increase with n ncrease n irspeed.

Examination f he lternating tresses n he otor- blade teel par t ach lade tation h o w s hat hey re all below he llowable lternating tress or nfinite fatigue ift f 25,000 ounds er quare nch.

♦

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BLADE FLAPPING MOTIONS

Blade f l app ing da ta are given In terms o f the s t eady f i r s t harmonic l o n g i t u d i n a l and la tera l f l a p p i n g componen t s wi th respec t to the shaf t . hese resu l t s fo r al l s teady f l igh t cond i t ions over the o p e r a t i n g speed range are presen ted In Figu re 155* he ver t i ca l spread In the t es t poin t s at a g iven speed Is predomina te ly due t o cyc l i c »t ick t r i m p o s i t i o n var ia t ions i n fu se l age angl? at tack . It Is evident that even with the changes In opera t ing cond i t i ons , the f l app ing angles are r e s t r i c t e d to r e l a t ive ly l imi ted ampl i t udes o t exceed ing 6 degrees l o n g i t u d i n a l l y o r 4 degrees l a t e ra l l y.

ROTOR PITCH LINK

The s teady r o to r p i t c h l ink loads were main ly c aircresslve an d l ow in magni tude .

The a l te rna t ing p i t c h l ink loads were sma l l at low airspeed an d inc reased gradua l ly an d s t ead i ly wi th speed.

In genera l , they were o f l e sse r magni tude than o r i g i n a l l y p r e d i c t e d and summar ized in PiAC Report 16-S-27 , where the maxi&us a l te rna t ing p i t c h l i nk l oad der ived was +562 pounds at 220 knots . l l par t s have been analyzed fo r This fa t igue load» and m in im u m fa t i gue l i fe of pa r t s under th is l oad i s ' 1 2 0 hours (see Figure 156) .

ROTOR SHAFT TORQUE

The otor haft orque s well elow he llowable imit torque f 111,703 pound-inches- ultimate orque 62,555 pound-inches) for which t as esigned. T he lot f otor shaft orque ersus irspeed Figure 57) h o ws he easured steady orque or oth evel light nd orward peed limb.

The lternating orque s ery mall nd ncreases slightly with n ncrease n irspeed. It s well elow the llowable esign

lternating orque f 19,523 ound-

inches or nfinite atigue ife.

R O T O R H A F T IFT

T he otor haft lift urve (Figure 58) hows, as x- pected, hat he teady ift n he otor ecreases with irspeed ecause f nloading y he wing.

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246

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the 1073-pound llowable lternating oad or nfinite

design atigue ife.

TABLE V I I I . F L I G H T AND BLIP NUMBERS FOR D A I A USED I N ROTOR B L A D E FLAPPING A N G L E CURVE ( F I G U R E 1 5 5 )

F L I G H T B L I P FLIGHT B L I P

219 5 357 1 , 2 , 3 , 4 , 5 , 6 | 27^ 1,2,3,5.6.7 392 9

285 2,3, ,5,6,7 394 2 , 3 , 4 , 5 , 6 , 7 , 8 , 9 290 5,6,7,9,10 396 1,2 292 1 0 , 1 1 , 1 2 397 1 , 2 , 3 , 4 , 6 3 0 0 t.s 398 4 , 5 , 6 , 7 308 4 . 5 . 6 399 4 , 5 , 6 , 7 , 8 , 9 , 1 0 3 1 0 1 , 2 11 , 1 2 , 1 3 3 1 2 4 400 4 , 5 , 6 , 7 3 1 8 3 404 2 , 3 , 4 , 5 , 6 , 7 3 3 2 1 , 2 , J , 1 » 405 1 .3 .4 333 1.8 ' , 0 6 1 .2 .3 350 1 , 2 . 3 , 4 , 5 , 6 , 7 407 2 , 3 , 4 3 5 5 4,5.6,7 409 1.3 3 5 6 2 , 3 , ^ , 5 , 6 , 7 . 8, 9 411 1 . 2 J

ROTOR-HEAD ORIZONTAL ORCE N D O M E N T N D EACTIONS T PPER A N D O W E R O T O R HAFT UPPORT EARIU(35 (See igure 159)

The ertical otor haft as loads pplied o t t he rotor head: the ift, a ocal oment caused y ffset f he rotor lade orizontal in rom he xis f otation, nd horizontal orce aused y rag n he otor nd nclination of he otor hrunt ector. From train gage measurements n the otor haft aken n light ests,the ending-iboment ia- gram, hear iagr»' reactions f he pper nd ower earings,

and orizontal orce t he otor ub ay e btained s ol-

lows .

AsTuming hat o ixity n ending s fforded y ither the pper r ower earing, s n he tress nalysis, he horizontal eaction n he lower earing an e btained directly from he oment at the lower ending age I

R. » h MLQ /21.87 (68)

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BÜÖTOR BRG. S H E A R M O M E N T

Fig u re 159 . Geometry of Rotor Shaf t (Showing Be a r i n g Supports, St r a in Gages for B e n d i n g and Ty p ica l Shear and Moment Diagrams with Sy m b o l s . )

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The lope f he endlng-m om ent urve s constant ntil It eaches the pper earing upport nd he alue f he bending oment t he pper earing s

M U B 1.5071 LQ (69)

Above he upper earing, the oment curve Is also a straight ine f onstant lope. Knowing he ending oment at the pper earing nd lso t the pper ending age, he

moment at he otor ub an e etermined. The otor ub moment s

MH U G [ UB U G) x llleij

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The loca l moment at t he r o t o r head Is caused by t he ef - fect ive offse t of the ro to r blade l i f t . A l l o the r moments are caused by drag or long i tud ina l o r l a t e ra l forces on the ro tor.

The ro to r l i f t lo known f rom f l igh t tes t measurements .

The hor izon ta l force of t he ro to r hub F^ may be o b t a i n e d f rom k n o wi n g the moment at t he upper bear ing :

' H (Mrm - Mu ) / 19 .^7 U B ' H (71)

The orizontal eaction t he pper earing ., an e

obtained y aking oments about he pper age ocation:

For heck,

Ru (H.60R L MUQ)/ 11.64 (72)

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Tabulations f hese loads and om ents on he rotor haft fol- low or arious flight conditons.

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R O T O R H A F T E N D I N G

The rotor haft lternating ending oments at he ower and pper train age ocations # he orizontal oad t he rotor head, he orizontal eactions t he ower nd pper bearings, and he oment t ne otor head re lotted ersus airspeed f he ircraft n evel light nd orward-speed climb (Figures 160 through 65).

They all xhibit imilar haracteristics eginning with a alue t over, hen ecreasing lightly with n ncrease n

airspeed .o bout 80 nots, nd inally ncreasing with urther ncrease n peed.

For irspeed p o pproximately 160 nots, the otor- shaft oments re within he atigue imits for nfinite ife. Moments t peed bove pproximately 160 nots could imit the ife f he otor shaft, epending n he esults f ench fatigue ests, hich er e ot ade nder his rogram,

TAIL R O P E L L E R HAFT

A lot f propeller-shaft orsion ersus forward peed s shown n igure 166, It h o w s hat he teady orsion n he

propeller haft ncreases with n ncrease n irspeed. T he alternating orsion emains constant with irspeed t alue of pproximately 1000 nch-pounds. This s well elow he fatigue esign llowable f 1000 il650 nch-pounds for nfinite atigue life.

TAIL R O P E L L E R

Propeller ub ending oments re ather mall. The steady oments ecrease lightly with n ncrease n irspeed, while he lternating ending moments emain early onstant with orward peed (see igure 167).

The ail ropeller tresses re calculated sing he bending om ents measured n light. Centrifugal-force tress is added irectly o he teady ending tress o btain otal steady tress. The calculated ending tresses re lightly conservative, since he inimum ection odulus is used. Actually, he irfo.il ection s nsymmetrical, o hat although he oment of nertia s one value for ny articular blade tation, the distance from he eutral xis to he x- tr em e iber iffers etween urfaces.

Propeller lapwise tresses t lade tations 13,2 (kO

percent) and 6.5 (50 ercent) show inimum f catter

(Figures 168 nd 69), ut for hose t lade tations 23.1

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airspeed Figure 7^).

here s pproximately he ame aisni- tude f lternating tress n ach ongeron^ o nly ne urve was drawn o epresent both.

The lternating tress was small nd ncreased lightly

with n ncrease in irspeed.

i . MAIN LAUDING GEAR

Axial load elt by the nock trut and xial load nd bending n he ain ember ere easured n run-on landing. A

plot of hese loads versus time for ypical un-on landing r

s hown n igure 176.

All oads are ell within he limit design loads for he gear. Maximum oad n he shock trut in his typical un-on

landing s 1900 ounds, compared o imit load or he hock strut of 560 ounds compression.

Maximum oad n he ain ember s 920 ounds compression plus 16,500 inch-pounds bending, hereas limit design load s 3830 ounds compression lus 6,700 nch-pounds moment. Other recorded oads in anding ere less han hese.

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T E C H N I C A L P R O B L E M S

»

D u r ing tne performance of th e ground- and f i ign t - t es t program with th e l 6 H - l A , a numoer of problems developed, most of which had th e effect of causing delays unt i l they were so lved ut di d no t prevent cont inuat ion of th e f l i gh t - test program. F or th e most part, the problems encountered were the mechanica l structura l type common i n the deve lopment of new equipment rather than "state-of-the-art" knowledge .

P R O B L E M S E N C O U N T E R E D D U R I N G G R O U N D T E S T I N G

These were p r ima r i l y structural i n nature an d were t yp ica l of th e tes t ing of almost an y new dynamic components.

1 . O i l pump d r i v e - O n two occasions, a draw-bolt w h i ch holds a l ube-o i l -pump d r i v e adapter i n d r i v i n g en- gagement w i t h a gear shaft i n the t ransfer gearbox f a i led at th e base of the threads. O ne of the f a i lu res occurred on the pressure-pump d r ive of the lube sys tem; th e o ther occurred on the scavenge-pump d r ive i n a lmost the iden t ica l manner. he f a i lu res were a t t r ibu ted to la tera l v ib r a t i o n of th e bol t , which , i f t i gh tened to the standard torque, w o u l d be nea r ly I n resonance with th e gear- tcoth-mesh f requency and could be brought exac t ly to resonance by over torquirg. T he des ign was mo d i f i ed to Incorporate a phenol ic p lug i n s ide the h o l lo w gear shaft whicn surroundad the draw bol t and prevented I t f rom v i b ra t i n g l a te ra l ly.

On two other occasions, th e lube-pressure-purap d r iv e shaft f a i led at a keyway. n both occasions, the pump continued to run, and th e f a i lu re was only found after d i sassembly for other reasons. he f i rs t f a i l u r e was at t r ibuted to fa t igue caused by inadequate chamfers on the key and Inadequate f i l l e t s i n the keyway. Afte r the second fa i lure , however, the key d r ive was rep laced w i th a sp l ined connect ion many t imes stronger. A f t e r

only 3 0 minutes of operat ion w i th the new s p l i n ed shaft, th e shaft suddenly se ized i n i ts journal bearings , shea r ing th e s p l i n ed sect ion and caus ing s toppage and se izure of th e oi l pump. A na lys i s of th is f a i l u r e gave pos i t i ve ev idence of the basic t rouble . Def lec t ions of th e shaft connect ion r e l a t i ve to the pump were causing a rota t ing bending moment between th e pump shaft and th e gear shaft to w hic h s t i f f sp l ine connect ion permit ted be c a r r ied in to the pump, causing bronze Journal bearing. Ex p an s i o n

i t was spl ined. T he the bending moment to loca l h ea t i n g of i ts

of th e pump-hous ing face eventua l ly caused a se izure w h i ch sheared the

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spline n ure orsion, (The am e -ctlon ad pparently previously roken ut he former keyway, ut hen he rigidity as reduced nough o hat he rive ontinued to unction with he roken iece rapped n lace.) At this oint, the um p rive as edesigned o ncorporate a loating quill haft, pllned t oth nds, hich could not ransmit oment esulting rom gear eflec- tions. The unp rive hen ave o urther rouble.

2. racks n he fuselage kin art of he round ests

Involved unning t igh ower evels n he otor nd propeller, with he ircraft ied ow n o he oncrete apron. As esult f ocal oad oncentrations t the ie-down oints, cracks hich equired epair e- veloped n he uselage kin nder hese oints. These loads er e aused y he ethod f estraint during he ie-down ests, nd er e ot epresentative of light conditions. Local atigue racks ecame evident n he airings nd n red ling-edge tructure. Sandwich anels er e ubstituted or he iveted airing skins which were eing oil anned" nd howing igns of racks. After uch modifications, o urther racks appeared or he emainder f he round esting r

throughout he light ro g ram.

3. ropeller pitch-control earing - After 8 ours of running, xcessive lay eveloped n he ropeller itch- control mechanism. Disassembly evealed hat he ronze retainer-cage ad roken n all earing nside he propeller upport ox. This as eplaced with tainless-steel age f heavier ross ection, hich ad to e upplied y he earing manufacturer. The earing caused o urther rouble fter his epair.

4. Propeller pinner Near he conclusion f he 0-hour

ground unning, the ropeller pinner uffered atigue failure t n ttachment oint which esulted n ts tearing oose. It as truck y he ropeller, disin- tegrated, and aused om e damage o he ropeller lades and he ing-tail uct. The amage o he laues and the uct as repaired romptly, ut he pinner equired redesign. Some f he arly light esting, involving hover nd ow-speed lights, as erforraed without spinner.

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P R O B L E M S E N C O U N T E R E D D U R I N G F L I G H T T E S T I N G

1. Turbine speed ont ol he -58 urbine s quipped with uel ontrol ystem which ncludes a peed- selector ever, the osition f which causes overnor to upply fuel n esponse o hanges in utput peed. The ange f he overnor etting, at o oad, la from about 50 0 to 80 0 R P M (at he utput of he peed- decreaser earbox). Because of built-in roop characteristics of he governor, this range, at full load,

is from 0 0 0 to bout 20 0 RPM. If he lever s moved beyond he lowest overnor etting, it asses through a ransition egion here he overnor will ot control the peed, thence o round dle, nd inally, to ut- off. Friction nd ther orces acting n he lever are uch hat ilot-effort oads on he onventional twist-grip control ere too igh for irect echanical linkage. Hence, a roduction lectric ervo ystem was nstalled wherein he osition f he wist-grip positioned potentiometer dualized or afety f flight), he ignal rom which, n urn, ositioned an electric ctuator onnected o he urbine oeed selector ever.

During he first utorotative descent n he 16H-1A the turbine speed-control ystem malfunctioned and the flight nded with n ctual mergency ower-off landing. When he ilot ttempted his lanned ower recovery y wisting is grip o he full-speed osition, the urbine did ot respond, ut emained t round idle. As a esult of anding n ough ield djacent to he airport, damage as ustained by he ail anding gear, ropeller, duct, and udder anes, ll f hich w ere epaired n ue ourse. The peed-control ailure was found o e an lectrical pen ircuit hich e-

veloped nside he ctuator.

To nsure that such an vent could ot reoccur, a limit-switch n he actuator as adjusted o hat the actuator ould ot run eyond he lower imit of he governed ange nless a eparate witch ere operr.ded from he ockpit. This switch as used nly to hut down he urbine, and was always ept,in he afety

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posi t ion I n f l ight . Thus , no elect r ica l fa i lure of any sort could place the speed se t t ing lower than the m i n i - mum governed set t ing, where near-ful l power would be Immed ia t e ly avai lab le . Th i s so lu t ion permit ted trouble- free operat ion for the remainder of the test program. However, i t di d place a res t ra in t on s im u la t in g abrupt power fai lures . he R P M of the autorotat lng rotor could not drop oelow that main ta ined by th e governor at i ts m i n i m u m set t ing, which , at no load, was only 3 percent below the normal powered R PM. A s soon as the decay ing R P M reached that value, the governor would

demand fuel and the tu rb ine wo u ld supply torque to sus ta in that R P M .

2 . Measurement of w i n g l i f t - O ne of th e object ives of the f l igh t test program was to eva lua te the effect of var ia t ions i n l i f t ' i s t r ibu t ion between w i n g and rotor on per fo rmance . I * . , vas planned to measure the rotor thrust by measur ing tension i n the rotor shaft , and to measure wing l i f t Independen t ly wi th s t ra in gages on t he fuse lage frames to wh ich the wing spars are attached. C al ib ra t ion of both systems was performed by hanging t he a i rc ra f t at the rotor, through a dynamometer. Read ings were t aken of the rotor-shaft t ens ion st rain gage b r idge and of the tw o br idges on each s ide of the fuse lage f rame which , when combined, measured the wing l i f t . The weight of the a i rcraf t was then incrementa l ly s u p p o r t e d on t he wings and unloaded from th e rotor w h i l e t he changes in read ings were recorded. hi s procedure was repea ted with two di fferen t spanwise points of suppor t fo r the w ings , to permit a so lu t i on independent of the spanwise center of l i f t , which was not known preci se ly. A ser ies of m u l t i p l y i n g constants for the four wi n g bridges was worked out which checked th e ca l ib r a t io n for both sets of loadings . owever, th e i n- f l igh t readings for these four bridges were so var i ab l e

and inconsis ten t that no

useful

data could be

der ived f rom them, e i ther i n d i v i d u a l l y or i n combinat ion. T he data from the rotor-shaft tens ion, i n conjunct ion wi th the known gross weights , on th e other hand, were suff ic ien t ly consis tent to supply the required informa- t ion. Consequen t ly, they were used I n this report to the ex c lu s io n of the w i n g l i f t measurements .

3 . Turb ine ai r in take - I n order to make use of ex i s t i ng gear ing and other power- t ransmiss ion components, the T- 5 8 tu rb ine was mounted i n a bur ied i n s t a l l a t i on i n s ide the fuselage, behind the rotor t ransmission. T he a i r- in take was designed to be of the s ide- in le t type, w i t h a forward-facing scoop. T he ai r entered a

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plenum chamber af ter making a 90-degree turn nd then entered th e turbine through a be l lmou th i n i e t from another ex i s t ing he l icop te r i n s t a l l a t io n . Ear ly i n th e f l i gh t test program i t became ev iden t that th e in let losses i n hover were so severe that the aircraf t performance was l i m i t e d by turbine i n l e t temperature at power l eve l s wel l below th e turb ine rat ing. T he losses were main ly caused by too sm a l l a rad ius i n the 90 - degree turn. T he external por t ion of th e i n l e t was therefore removed, so tuat th e ai r s im p ly entered a rectangular f lush opening i n th e si Je of the fuselage, th e area of wh ich was over tw ice as l a rge as the throat of the or ig ina l external in le t . Although t h i s expe- diency di d not permit any ram-pressure recovery, i t reduced th e i n l e t losses su ff i c ien t ly that most of th e f l igh t test program was conducted wi th th i s conf igura- t ion. F or f l igh ts at speeds above about l 6 0 knots, however, the problem of l i m i t i n g turbine i n l e t temperature again arose. B e g i n n i n g with test 403 , the ex - ternal i n l e t was r e ins ta l l ed i n conjunct ion with an improved be l lmou th f rom the p lenum chamber i n to the turbine. A n improvement i n turb ine performance at h igh speed was noted, but l i m i t turb ine i n l e t temperature was always reached before rated power could be

obtained. A n Improved i n l e t was des igned and fabr ica- t ion was started, but i t was not completed i n t ime for use i n th e test program. I ts use would h av e increased th e power at ta inable at high speed by an es t imated l 60 horsepower, wh ich would have increased th e true airspeed to 17^ knots.

4 . Rotor control sys tem s t i ffness - T he 1 6 H - 1 A , as modi f i ed from th e 1 6 H - 1 and as ground tested, used an H - 2 1 rotor, I n c l u d i n g blades, hub, and swash plate. T he remainder of th e rotor control sys tem was the same as that f lown i n th e 16H-1 , w i t h only those changes necessary to adapt to the H - 2 1 swash plate. T he system i n- corporated three i r r ever s ib le hydrau l ic servo actuators , one each for longi tudinal cyc l ic , la teral cyc l ic , and co l l ec t iv e pi tch. T he system was s t ress -checked for th e an t ic ipated h igher loads of th e larger rotor and h igher f l i gh t speeds and was found to be adequate.

I n th e course of exp lo r ing progres s ive ly h ighe r f l i gh t speeds, a v ib ra t ion appeared at about 120 knots which increased rap id ly wi th speed and became severe enough to l i m i t the f l i gh t tes t ing to about 140 knots. I n- v es t ig a t io n of the f requencies of v ib ra t ion showed the largest contr ibutors to be the second, th i rd , and tenth harmonics of rotor speed. T he tenth harmonic was

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a t t r ibu ted to the ta l i p rope l le r, which has a speed o f 9 -7 t imes ro to r speed. Inspec t ion uncovered a malad jus tment which resu l ted in one p r o p e l l e r b lade having a p i tch of near ly 2 degrees more taan t he other two. Correc t ion of this cond i t ion e l imina ted most of the t en -pe r- r evo iu t ion v ib ra t ion . Attempts were made to reduce t he th i rd harmonic componen t with a tuned dynamic v ib ra t ion absorber which was placed at var ious loca t ions in t he a i rcraf t rang ing f rom the center o f grav i ty to a pos i t i on Just ahead of the cockp i t nose. Figure 177 shows the effec t of the absorber mounted in th is las t pos i t ion , an equ iva len t

dead weight mounted there, and a l so with noth ing mounted. Although t he th i rd harmonic componen t was s ig n i f i c an t ly reduced , t he other vibra t ions not ab- sorbed resu l ted in a compos i te vibra t ion spec t rum s t i l l u nsa t i s f ac to ry t o the pi lo t .

At th is po in t , a mot ion pic tu re camera was i n s t a l l ed on the py lon to take in - f l igh t p ic tu res of swash-p la te mot ion . he resu l t ing pic tu res c lear ly showed swash- p la te motions a l lowing the ro to r blades ' two- and th ree- p e r- r evo lu t ion p i tch changes of t he order of +.5 degree , a l though the contro ls in t he cockp i t showed prac t ica l ly

zero mot ion . he s t i ffness o f the con t ro l sys tem cables , pu l ley brackets , push-pu l l rods, and be l lc ranks between t he swash p la te and t he i r revers ib le ac tua tors under t he f loo r was insuff ic ien t , and the r o t o r blades were changing pi tch in r e sponse to harmonic ai r load componen t s .

To correc t the s i tua t ion , the ro to r con t ro l sys tem was changed to incorpora te a set if dual i r revers ib le hydrau l ic servo ac tua to rs , mounted in t he py lon immedia te ly under the swash pla te , in the p lace of those under the f loor. Thus, mot ion of the swash p ia te was ar res ted s t i ff ly at i t s source . This change, which

lowered t he v ib ra t ion level markedly, permi t ted tes t f ly ing to con t inue and to proceed to the maximum speed reached in the p rog ram (195 knot s ) . Figures 3 8 th rough ' » 7 show the Improvement th roughou t the speed range brought about by the con t ro l s t i ffen ing .

5 . Prope l le r blade des ign - Measurements of t a i l prope l le r power at speeds up to 140 knot s ( the temporary vibra- t ion l imi t , see Roto r cont ro l sys tem s t i f fne s s discussed above) ndica ted that the prope l l e r-b lade ang le o f a t tack at h igh speed wo u l d be above op t imum, w i th a t tendan t low prope l le r e ff ic iency. During t he per iod in

which the con t ro l sys t em was be ing modif ied , new blades

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were fabr icated for the propel ler. A l l of the data preiented for speeds above 1 ^ 0 knots was obtained wi th th t new p ro p e l l e r blades . Data obtained i n hover and at e l ower forward speeds wi th th e new p ro p e l l e r were not s ig n i f i can t ly different from those obtained wi th the or ig ina l one.

Pod-end i n rotor control sys tem - Dur ing th e period of v ib ra t io n inves t iga t ion , pr ior to control s t i f fen ing , one of th e rod-ends i n the long i tud ina l cyc l i c control system f a i l ed at the bear ing-re ta in ing lug. Upon ro l - l i n g out of a c l i m b i n g turn , there was a loud noise and the ai rcraf t pi tched to a nose-high atti tude. Control was recovered, and a caut ious check revea led at leas t part ial control about al l axes. he p i lo t dec ided to make a ro l l - o n l and ing s ing propel le r pi tch control to control rate of descent. Dur ing the approach and l anding , the c y c l i c s t ick w as used only for lateral control . Long i tud ina l t r im was main ta ined wi th th e e lev a to r t r im, and co l l ec t iv e pi tch was not used.

normal ro l l - o n land ing was performed, wi th touchdown between 5 0 and 60 knots.

A l l l o n g i tu d in a l , l a tera l , and co l l ec t iv e push-pull rods, bel lc ranks , levers , and quadrants subject to rotor control loads were inspected, w i th no other ev idence of damage. owever, th e fa i l ed rod-end and others s i m i l a r to i t, i f subject to loads of th e same order, were rep laced wi th new, stronger parts.

T he a b i l i t y of the e leva tor t r im to serve as a backup for l o n g i tu d in a l cyc l i c pi tch i n an emergency was a feature not apprec ia ted i n th e or ig inal design concept I ts emergency usefulness proved to be a s ign i f ican t asset, and i n this ins tance i t probably saved the aircraft .

Landing-gear emergency ex tension - T he l and ing gear

i s extended m an u a l ly on the 1 6 H - 1 A i n th e event of fa i lure of the normal e l ec t r i ca l ly operated retract ion system. he emergency sys tem consis ts of a hand- crank i n the cockpi t , on which i s mounted a sprocket connected by ro l l e r- ch a in to another sprocket on the retract ion actuat ing shafi . he hand-crank dr ives through a spr ing- loaded clutch, so that i t remains stat ionary dur ing e lec t r ica l actuat ion of th e gear, bu t th e chain and sprockets are i n motion. O n on e occasion, w h i l e th e aircraft was i n th e shop for maintenance, the r o l l e r- ch a in broke during a t r ia l gear- re t ract ion cycle and became Jammed, preven t ing

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both e lec t r ica l and manual operat ion.

T he f a i l u re was traced to th e faulty i n s t a l l a t io n of th e removable s i d e - l i n k at th e spot where the two ends of chain ar e jo ined to make the chain "endless". A f a i l u re of t h i s type occurr ing i n f l igh t wo u ld h av e prevented a . normal landing. Hence, to preclude such an event , the sprocket on th e m ain actuat ing shaft was remounted to d r i v e through a shear-pin , designed to shear at a load wel l above the hand-crank ex tens ion load bu t we l l below th e torque of th e e lec t r ic actuator. T n i s would permit e lec t r ica l operat ion, even wi th a Jammed chain. N o fur ther d i f f i cu l ty was encountered

wi th th e

re t rac t ion system.

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C O N C L U S I O N S

i . The use of a v a r i ab le pi tch pusher propel le r mounted I n- s ide a duct, i n conjunction wi th m o v ab le d i rec t iona l and long i tud ina l control surfaces, prov ides a convenient and eff ic ien t method of effec t ing th e desired power d i s t r i bu - t ion between that u t i l i z e d for forward propuls ion and that required for rotor l i f t over th e entire speed range. Power management can be accompl ished eas i l y by control of both rotor and propel le r pi tch by on e p i lo t , w i t h on e hand.

2 . T he use of a ducted pusher p ro p e l l e r for forward propul- s ion of a compound hel icop ter i s an effec t ive method of ex ten d in g the speed capab i l i t y by r e l i e v i n g the m ain rotor of i ts p ro p u l s iv e function. At speeds above 150 knots, th e rotor required as l i t t l e as 15 percent of th e total power used.

3 . T he stators and the propel le r duct render a h igh degree of long i tud ina l and direc t iona l s ta t i c s t ab i l i t y and produce aerodynamic damping for dynamic s tab i l i ty. he longl tud inal control surface acts as an effec t ive t r im control even i n a hover, where i t produces an appreciab le moment

equiva len t to a center-of-gravity sh i f t of ' l .1

» inches i n the 16H - 1A .

k. T he employment of a ducted pusher p ro p e l l e r and of de- f lector vanes i n a compound he l icop te r i s a feas ib le and adequate means of main- ro to r torque compensation and di - rec t ional control . Control was found to be suff ic ien t for h o v e r in g and maneuver ing s ideward 0 to 30 knot« an d rearward 0 to 30 knots .

5 . T he employment of a ducted pusher propel le r w i t h rudder control surfaces provides favorab le direct ional control for making running ( S TO L ) takeoffs. Takeoffs at ground speed as h igh as 5 5 knots were conducted dur ing t h i s test program.

6 . I n the conf igura t ion tested, th e 1 6 H - 1 A was power l imi ted . L e v e l - f l i g h t true airspeeds approaching 20 0 knots wo u ld be poss ib le wi th an engine a i r- in take and wi th th e incorporat ion of drag-reduction i tems.

7 . T he m ax im u m true airspeed of 1 95 knots was a t ta ined i n a 10-degree dive . T he l i m i t i n g v ib ra t io n l ev e l s , exper i - enced i n the ear ly part of th e test program i n the speed range from 11 0 to l ^O knots, were a l l ev ia t ed by the i n-

corporat ion of a s t i f fe r rotor control system.

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T he steady component of the rotor-blade bending moments and p i t ch l i nk loads, and th e a l te rna t ing component of propel ler-b lade bending moments and p i t ch l i n k loads, tend to remain nearly Independent of ai rspeed. O n the other hand, the a l t e rna t i ng component of rotor-blade bend- i ng moments and p i t ch l i nk loads, and the steady component of propel ler-b lade bending moments and p i t ch l i n k loads, tend to increase gradual ly w i t h ai rspeed. o wev e r, th e net fat igue effect of th e combined l oad ings remained w i t h i n ample st ructural margins and showed only a gradual Increase wi th speed.

Si n ce th e w i n g supports an ever greater proport ion of th e a i rcraf t weigh t as speed increases , and the rotor supports a correspondingly smal l e r propor t ion , th e rotor at some point begins to lose I ts effec t iveness for con- t ro l l i ng th e a i rcraf t i n ro l l . F or compound ai rcraf t des igned to unload the rotor much beyond 5 0 percent. Incorporat ion of convent ional a i l erons should be bene- f ic ia l i n two respects.

a . R o l l control at h i g h speed should be improved.

b . Si n ce the rotor now need not supply al l of the roll control at h igh speed, i ts la teral cy c l i c control can be reduced wi th r esu l t i ng improved hover ing ro l l response.

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RECOMMENDATIONS

Plight tes ts of the r ing- ta l l sha f t -d r iven compound h e l i - copter should be continued with a f l l ghc test a r t ic le hav ing add i t i ona l powt r and an improved ai r i nduc t ion sys t em to permit exp lo ra t ion of th e f l i gh t cha rac te r i s t ic s at higher speeds.

The a i rcraf t for these extended tests should incorpora te a i le rons wi th an adjustable l inkage ra t io i n the la tera l cycl ic cont ro l . The f l ight p r o g r a m should Inves t iga te var ious p r o p o r t i o n s o f these t wo la tera l control inputs to obta in the opt imum con t ro l s for both low-speed, he l i - copter type f l igh t and high- speed unloaded rotor f l ight .

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R E F E R E N C E S 1 . oldberg, J . H . , Automat ic Control : Pr in c ip l e s of Sy s tem s

Damamlcs . A l l y n and Bacon Incorporated, Boston, Kassa- chuset ts , 1961 .

2 . eckel , E . , S t a b i l i t y an d Control of Airp lan es and Hel i co p te r s . Academic Press , N ew York, 19&1

3 . -2 1 C Phase I V - Performance Ev a lu a t io n , Edwards A i r Force F l i g h t Test Center, Cal i fo rn ia , Report A F F T C - T R - 5 5 - 2 6 , December 1955 -

^ 4 . Hel icop te r F l y i n g and Ground H a n d l i n g Qu a l i t i e s ; Ground Requirements for . M i l i t a r y Spec i f i ca t ion M I L - H - 8 5 Ö 1 A , September 1961.

5 . e l i cop te r Performance Tes t in g Manual. Nava l A i r Tes t Center Test P i l o t S choo l , Patuxen t Nava l A i r Sta t ion , Maryland , January 1965 .

6 . hase I V Performance and S t a b i l i t y Tests of th e H - 2 1 Bf Edwards A i r Force F l i g h t Test Center, Cal i fo rn ia , Report A F F T C - T R - 5 7 - ^ , March 1957 .

7 . tapleford, R . L . , et al . Systems Techno logy, Incorporated, A n An a ly t i ca l S tudy of V / S T O L H a n d l i n g Q u a l i t i e s i n ;

H o v e r and Tr a n s i t i o n . Report A F F D L - T R - 6 5 - 7 3 . A i r Force F l i g h t Dynamics Laboratory, Wr igh t - Patterson A i r Force Base, O h i o ; October 1965 -

8 . apscott , R . J. , Cr i t e r i a for Control and Response Charac te r i s t ics of Hel i co p te r s and V T O L Aircraf t i n Ho v er in g and Lo w-Sp eed F l i g h t , Aerospace Engineer ing . Vo lu m e 19, June i 960 .

9 . eat Requirements. Ground. Hel icop ter. M i l i t a r y

Spec i f i ca t ion M I L - H - 8 6 7 9 , 5 March 1 9 5 ^ . 10. B u l l e t i n A N C - l B , Des ign of Wood Aircraf t Structures.

Aircraf t Committee, Mu n i t io n s Board, June 1951. Super in tenden t of Documents, U . S . Government Pr in t in g Off i ce , Washing ton , D.C.

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A P P E N D I X I

D E S C R I P T I O N A N D R E S U LT S O F G R O U N D T E S T S

1 6 H - 1 A W I N G P R O O F L O A D T E S T

T he purpose of th e test was to prove that the 1 6 H - 1 A w i n g could safely withs tand 120 percent of th e l i m i t loads expected I n f l i gh t tes t ing of th e model 1 6 H - 1 A aircraf t . T he test was performed w i t h th e w i n g i n s t a l l ed on th e aircraft.

T he w i n g was loaded downward i n s t ead of upward täte the tes t ing, al though th e greatest a i r load act Th i s was poss ib le , s ince phys i ca l l y the outboard wi i s symmetr ica l about th e horizontal Jhord l ine , an d sect ion i s nearly symmet r i ca l . w o l oad ing condit i executed. I n C ond i t i on I , the w i n g was un i fo rmly 1 spanwise and chordwise w i t h th e center of pressure 25-percent chordl ine, represent ing a maximum spanwi condit ion. I n C ond i t i on I I , the w i n g was un i fo rmly spanwise and chordwise w i t h the center of pressure 35-percent chordl ine. he l oad ing was 71 percent o l oad ing of C ond i t i on I , and represented a m a x i m u m w s ion condi t ion.

to f ac i l l - s uoward. ng panel the center on s were oaded at the se bending

loaded at th e f the i ng tor-

I n each of th e above l oad ing condi t ions , th e port w i n g was loaded to 1 2 C percent of th e expected f l igh t test load, w h i l e the starboard w i n g was loaded to approx imate ly 70 percent of th i s load. T h i s dissymmetry of l oad ing represents the unsymmetr ica l f l igh t cond i t ion and i s cr i t i ca l for the w i n g spar webs i n the w i n g center section.

A f t e r th e comple t ion cf th e test, the w i n g was inspected and was found to be free of an y ev idence of damage.

C O N T R O L S Y S T E M P R O O F T E S T S

T he purpose of these tests was to proof load the 1 6 H - 1 A control sys tem ( long i tud ina l and la teral cyc l ic , co l l ec t ive , and rudder) i n order to insure i ts structural in tegri ty and to determine whether an y exces s ive deflect ions ex i s t ed throughout the system.

F or tests beyond th e stops (tests 1 and 3) , th e rotor was locked out at th e main rotor pi tch locks on th e rotor hub, thereby prevent ing an y movement of th e swash pla te i n c yc l i c and co l l ec t i v e pitch. In al l tests, loads were app l i ed w i t h a tens ion scale to th e handgr ip of th e par t icu lar s t ick being tested. Boost pressure of 1 7 0 0 pounds per square inch was

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appl ied with hand-operated h y d rau l i c equipment. F or th e test up to the stops (test 3), th e fore and af t loads on the cyc- l i c s t ick were ap p l i ed wi th the ai d of a tu rnbuck le I n ser ies wi th a spr ing balance. T he system was observed throughout, and def lec t ions of p i lo t s t ick were measured.

T he rudder was blocked out at th e ta l l w i t h a wood frame. Boost pressure was ap p l i ed w i t h hand-operated hydrau l ic equipment. Proof loads were ap p l i ed to the foot pedal by the test operator s i t t i n g i n the p i lo t ' s seat. Loads were measured wi th a compression scale.

Resul t s

T he cyc l i c and c o l l e c t i v e pi tch system withs tood al l loads sa t i s fac to r i ly, and no apparent ex cess iv e def lect ions were noted i n th e p u l l ey s or brackets i n tests 1 and 2 . I n test 3 , def lec t ions and permanent se t occurred i n the stop of th e copi lo t ' s s t ick and were corrected by s tructural modif ica t ion .

T he rudder sys tem withs tood al l loads sa t i s fac to r i ly. T he on ly apparent def lec t ion not iceab le was the "walk ing beam" which , under r igh t pedal load, def lec ted a s l i gh t amount.

A I R F R A M E S H A K E T E S T

T he purpose of th is test was to determine the magni tude of v ib ra t io n throughout the aircraf t w h i l e i t waa suspended from the boom and t i ed down i n the ground tes t ing i n order that any cond i t ion considered undesi rab le , in to le rab le , or unsafe could be corrected before f l igh t tests. I t was no t Intended to supply absolute v ib ra t io n l eve l s , as the en- vironmental condit ions of these tests wo u ld no t permi t the s im u la t io n of th e same v ib ra t io n l eve l s that w i l l occur i n f l ight . Vib ra t io n l eve l s var ied great ly w i t h th e tens ion on th e t i e -down cables , th e number of t ie -down cables that were Ins ta l led , the dead load placed i n the aircraf t , and the degree of cyc l i c s t ick used. A l s o , ground win d condi t ions affected the v ib r a t io n l eve l s .

T he most severe v ib r a t io n s were obtained w i t h th e a i r- craft t i ed down and occurred i n the rudders, hor izonta l sta- b i l i ze r, an d stators. Some v ib ra t io n also occurred i n the

♦ s ide sk in of the fuselage. ib ra t ion i n th e cockpit was minor. Some of th e v ib r a t io n that th e aircraf t exper ienced w h i l e t i ed down was Induced or aggravated by the t i e -down cables. T he a ircraf t , when undergoing tes ts suspended by i ts rotor head, was observed to ru n wi th less v ib ra t ion than

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when t i ed down to the ground.

Exces s ive v ib ra t i on I n the rudder control cable I n th e bot tom hatch area was observed. A l s o , s l i g h t ver t ica l move- ments of th e rotor l ong i t ud ina l control rods a t th e swash plate were noted. I n order to improve maintenance i n the areas ou t l i ned above, th e f o l l o w i n g was accomplished:

T he rudders and f i t t ings were s t i ffened an d the toler- ances i n the rudder controls were reduced to a m in im u m to reduce dynamic loads on the system. T he hor izonta l s t ab i l i - zer was reinforced by add i t i on of doublers and sk in s t i ffen- ers , making use of bonding to increase fat igue l i fe further. T he t h in a lu m in u m sk in of the stators was changed to sand- w i ch cons t ruct ion .25 inch t h i ck w i t h a luminum honeycomb cores and a lu m in u m faces bonded together. he s ides of the fuselage, wh erev e r the th in a lu m in u m sk in was buckl ing , were re in fo rced wi th a . 6 2 -b y - .6 2 -b y - .0 3 2 - in ch 2 0 2 ^ T 3 angle s t i f - fener i n s t a l l ed c i rcumferen t ia l ly to break the 7 - inch sk in panels i n to two 3*5- inch -wide panels wherever needed. A n a luminum bracket w i t h a phenol ic guide was i n s t a l l ed to d i v i d e the rudder cable length i n two parts where i t was v ib ra t i ng excess ive ly.

T he ro to r long i tud ina l control rods and f i t t i ngs were s t reng thened*as well as the at tachment of the f i t t i ngs to the rods .

Changes that were made to the t ie -down arrangement were a s fo l lows :

1 . he ve r t i ca l rest raints , from th e pylon to th e ground, were moved to a more near ly ver t ica l pos i t i on i n order t o react the ver t ica l thrust loads as close to th e or ig in as poss ible .

2 . dd i t i ona l ve r t i ca l cables were added from the wing- root to th e ground i n order to take ou t th e hor i - zontal thrust load produced by the ro-or.

3 . orizontal rest raints were added, from eh e underside of the fuselage, near the center of gravi ty af t ap - prox imate ly 8 feet, i n order to take ou t the hor i - zontal thrust load produced by th e propel ler.

After th e incorporat ion of the changes described, th e v ib ra t i on l eve l was considered to be acceptable by th e p i l o t even a t th e extreme propel le r and/or co l l ec t i ve p i tch set- t ings. he degree of v ib ra t i on va r i ed from pract ica l ly neg-

l i g i b l e wi th controls neutral to moderate when the cyc l i c

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was f u l l y disp laced (near ly to th e stops). A t no t ime during th e

a s feedback, i n the c yc l i c though h igh frequency v i b r rudder pedals , no fur ther changes were incorporated. d u c t . o n f v i b ra t i o n caus a r r a n g e . - n e n t and the practi with th e aircraft suspende correlat ion between vibra t that m i g h ' : become apparent

T I E - D O W N T E S T S

t es t ing was any v i b ra t i o n fe l t , or co l l ec t i ve pi tch levers . A i -

at icns were i n i t i a l l y noted i n the v i b ra t i o n s were noted after the

Because of the s ign i f i can t re- ed by th e changes i n th e t i e -aown ca l ly v ib ra t i on - f ree ta i l tests d, i t was concluded that no i on dur ing ground tests and those

during f l i gh t test could be made.

T he purpose of these tests was to subject the a i r f r ame systems, t ransmiss ion sys tem,and other dynamic components to a ser ies of tests to check t he i r st ructural in tegr i ty and funct ional operat ion pr io r to f l igh t . T he ground tests commenced 8 A pr i l 1965 and 45 hours were completed 8 J u l y 1965 . hey were conducted u t i l i z i n g the ent i re ai r f rame and dynamic components, i nc lud ing the rotor and t a i l propeller. Consequen t ly, the ai rcraf t had to be res t ra ined to resist al l loads induced by the mai n rotor, ta l l p rope l - ler, and rudder vanes.

T he f i rs t 15 hours of tes t ing were conducted pr imar i l y to prove the mai n rotor and th e t ransfer t ransmiss ion and to d i scover any discrepancies so that they mi g h t be corrected before f l ight .

T he main rotor loads (rotor l i f t and s ide loads) were taken out by four cables extending down at 45 degrees from the rotor py lon to anchor points I r the ground. hese cables were tens loned to 80 0 pounds, g i v i n g a total of l 6 0 0 pounds downward load. T o decrease the load i n th e cables dur ing high co l l e c t i v e sett ing, approxiraatley 1 0 0 0 pounds of bal l as t was pos i t ioned i n the ai rcraf t di rect ly beneath the

t ransmiss ion. Cables were a l so at tached to the w i n g f i t - t ings and extended to anchor points i n the ground to ac t as backup to the cables attached to the rotor pylon. T he thrust loads from th e propel ler were taken ou t by cables at tached to the lower longerons at s ta t ion 1 69 and extending af t at 25 degrees to the hor izonta l to anchor points on th e ground. T he side loads at th e ta i l were taken ou t by cables wrapped around s t a t ion 3^5 and extending l a t e ra l l y and downward to th e ground. T he ai rcraf t , i t se l f , was res t ing on wooden blocks w i t h the mai n and ta i l gear o leos comple te ly de- pressed to obv ia te an y poss ib i l i t y of ground resonance and to make the t i e -down conf igura t ion more r ig id . T he power

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M«» '

d is t r ibu t ion schedule shown I n Tab l e X V I was repeated te n t laes to complete the 15 hours.

TABLE XVI . T I E - D O W N T E S T P O W E R D I S T R I B U T I O N S C H E D U L E

TRANSMISSION H P T O H P T O T I M E PEST NUMBER M A I N R O T O R TA I L R O T O R ( M I N U T E S )

1 5 0 200 5 2 5 0 t o o 5 3 5 0 600 5

4 50 800 5 5 5 0 1 0 0 0 2 0 6 5 0 11 0 0 5 7 M00 200 5 8 HOO t o o 5 9 t o o 600 5

1 0 t o o 8 0 0 5 1 1 60 0 20 0 5 1 2 60 0 t o o 5 13 600 600 5 I t 800 20 0 5 1 5 800 t o o 5

T he test revealed a def ic iency i n the t ransmiss ion oi l pump bo l t ( 1 6 D 5 2 5 1 ) , the des ign of w h i ch was subsequent ly changed.

T he next 30 hours of tes t ing was performed accord ing to a schedule der ived rom paragraph 3.6 .3 3 of M I L - H - 8 6 7 9 (Re- ference 9), except that the t imes for uach test were reduced. D u r i n g the tests , data were recorded 0 1 rotor and propel ler stresses; rotor, prope l l e r, and tu rb ine power and R P M ; engine mount v ib ra t ions ; rotor and propel ler pi tch set t ings ; and powerplant and t ransmiss ion oi l temperatures and pressures.

R E S U LT S

1 . Transmis s ion Sy s tem

Wi t h the except ion of several o l ] -pump-dr ive fai lures w h i ch « re d i scussed under "Tecl nical Prob lems" , the t rans- m i s s i o n funct ioned sa t i s fac to r i ly. T he pump dr ives were redesigned before f l igh t t es t ing began, were retested, and gave no further trouble.

2 . Rotor S ys t em

T he rotor sys tem. I t se l f , performed without showing any

problems. However, shortcomings i n the ins t rumenta t ion came

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to l igh t during these tests. he rotor-shaf t- torque s t ra in gage bridges had two ac t ive arms on the shaft and tw o dummy arms i n th e external circui t . he br idges picked up spur i - ous s ignals from rotor-shaft bending which were super imposed on th e torque s igna l s and resul ted i n a loss of accuracy. T he br idges were rep laced wi th four-ac t ive-arm br idges for the f l igh t test program; these bridges proved much more sat is factory.

3 . Ta i l P rope l l e r

T he propel ler blade, hub, and pi tch l i n k st resses were wel l w i t h i n safe l e ve l s throughout the ground tests. I n- st rumentat ion operated sa t i s f ac to r i l y except for b lad e f l app ing moment at s t a t ion 29 .7 . everal at tempts were made to repai r th e gages at th i s s tat ion, but none w as successful . (They were f i n a l l y repai red pr ior to the f l igh t test phase).

A t the conclus ion of the 45 hours of t ie -down t es t ing , th e prope l l e r- r i ng - t a i l combinat ion was tes ted for an addi t ional 5 .8 hours. ur ing these ser ies of tests , the rotor blades and hub were removed and the a i rcraf t w as suspended by cable from a boom to the rotor pylon. upport- i ng structure at the r ing- ta i l was instrumented wi th s t ra in

gages to provide longi tudinal , la teral , and ver t ica l force data. rope l l e r torque, R P M , and pi tch , as wel l as rudder def lec t ions ere measured i n the normal manner y u s i ng th e 1 6 H - 1 A inst rumentat icn.

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A P P EN D I X I I P E RT I N E N T D ATA O N I N S T R U M E N TAT I O N , A I R C R A F T

D R A G , W E I G H T, A N D C E N T E R O F G R AV I T Y

Infttrumtntation used i n the f l igh t test program i s shown i n Tab l e X V I I .

TA B L E X V I I . l 6 U - i A I N S T R U M E N TAT I O N

t R A M E T E R

Mrspeed

Altitude

Clock

3as enerator

tngine R P M

Engine orque

Air Töqpera-

ture

Fuel low

Flight ounter

Vertical Rate

Rudder Po s i - t i on

Prope l le r

Pitch

C y c l i c Force Pi tch

C y c l i c Force Ro l l

"adder Force

R E A D O U T

Photo Panel

Photo Panel

Photo Panel

Photo Pane l

Photo Panel

Photo Panel

Photo Panel

Photo Panel

Photo Panel

Photo Pane l

Osc i l lo g rap h

Osc i l lo g rap h

Osc i l lo g rap h

Osc i l lo g rap h

Osc i l lo g rap h

T R A N S D U C E R

r . ' ; Reqd.

N ot Reqd,

N ot Reqd.

A.C. Generator

A.C. Generator

Autosyn

Resis tance

Autosyn

N ot Reqd.

N ot Reqd.

Potent iometer

S t ra in Gage

St ra in Gage

S t ra in Gage

S E N S I T I V I T Y C H E C K

N oc Reqd,

N ot Reqü ,

N ot Reqd.

N ot Reqd.

N ot Reqd.

N ot Reqd.

N ot Reqd.

N ot Reqcu

N ot Reqd.

N ot Reqd.

Fu l l Throw

Potentiometer F u l l Throw

Real»

Real»

Heal*

* S ee Note at E nd of Table .

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TABLE X V I I . Continued

PA R A M E T E R READOUT TRANSDUCER SENSITIVITY C H E C K

C o l l e c t i v e Oscillograph Strain Ga g e R e a l » Fo r c e

C y c l i c P i t c h P o s i t i o n

Oscillograph Fotentiometer F u l l Throw

C y c l i c R o l l Position

Oscillograph Potentiometer F u l l Threw

C o l l e c t i v e S t i c k Position

Oscillograph Potentiometer F u l l Throw

R u d d e r P e d a l Position

Oscillograph Potentiometer F u l l Throw

P i t c h R a t e Oscillograph Potentiometer R e a l »

R o l l Rate Oscillograph Potentiometer R e a l »

Y aw R ' . . f c j Oscillograph Potentiometer R e a l »

A n g l e of Attack

Oscillograph Autosyn Regulated P o w e r Supply

S i d e s l i p P h o t o P a n e l D . C . Äutosyn N ot R e q d .

R o l l A n g l e Oscillograph Potentiometer R e a l »

P i t c h A n g l e Oscillograph Potentiometer R e a l »

Rotor Azimuth Indicator

Oscillograph Magnetic P i c k u p

N ot R e q d .

B l a d e F l a p A n g l e

Oscillograph A u t o s y n Regulated Power Supply

B l a d e Lead-Lag A n g l e

Oscillograph Autosyn Regulated Power S u p p l y

B l a d e P i t c h A n g le

Oscillogra.h

c To f Ta b l e .

A u t o s y n Regulated Power Supply

» S ee No t e at E n

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1

300

TABLE VII. - ontinued

» A R A M E T E R R E A D O U T- TRANSDUCER SENSITIVITY CHECK

»Itch ink Load

Oscillograph Strain age Real»

ilade lap

Bending Station 6

Oscillograph Strain age Real*

Slade lap Bending

Station .5


älade lap Bending Station 9.2


älade lap Bending

Station 05.6


dlade lap Bending Station 21


31ade lap Bending Station 32


3lade lap Bending Station 58.4


älade lap Bending Station 84e8





Oscillograph

d f able,

Scrain age Real»

• ee ote t n

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TABLE VII. ontinued

SENSITIVITY CHECK

Real»

PA RA METER R E A D O U T TRANSDUCER

I 3lade Chord

Bending Station 46

Oscillograph Strain age

• •

Blade Chord Bending

Station 31.5

Oscillograph Strain age Real 1

Blade Trail- ing dge Station 2


Blade Trail- ing dge

Station 31.5

Oscillograph Strain age Real«

Blade Trail- ing dge

Station 10


Rotor haft Bending Upper 0 °


Rotor haft Binding

Upper °

Oscillograph Strain Gage Real»

Rotor haft Torsion


j Rote- haft Bending Lower 0°


Rotor haft Bending Lower °

Oscillograph Strain age Peal*

• Rotor haft

Lift Oscillograph Strain age Real»

» ee ote at End f able.

30 1

mmü**&timBflm*im**** i*y**w.-f**- - - -

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TA B L E X V I I , - Cont inued


Rea l »

PARAMETER R E A D O U T T R A i l S D U C E R

S t ra in Gage rop H ub Fl ap Bend ing

Osc i l l og raph

Prop Blade Flap Bending Sta t i o n 13.2

Osc i l l og raph St ra in Gage Rea l »

Prop Blade Flap Bending Sta t i o n 16 .5

Osc i l l og raph S t ra in Gage Rea l »

Prop Blade Flap B end ing Sta t i o n 2 3 . 1

Osc i l l og raph St r a i n Gage Rea l »

Prop Blade Flap B end ing Sta t i o n 2 9 .7

Osc i l l og raph St ra in Gage Rea l »

Prop Blade C h o r d Bending S ta t ion 13.2


Prop Shaf t Tors ion

Osc i l lo g rap h S t ra in Gage Rea l »

Prop Pi t ch Cont ro l Lever Load

Cockpi t St ra in Gage Rea l »

Prop Pi tch Link Load


Left Wi n g L i f t Forward


Left Wi n g L i f t A ft


Ri g h t Wi n g L i f t Forward

Osc i l l og raph St r a i n Gage Rs a l »

• See Note at E nd of Table.

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TA B L E X V I I . - Cont inued


PA R A M E T E R R E A D O U T T R A i J S D U C E H

Rig h t Wi n g Li f t A ft

Osc i l l og raph S t ra in Gage R e a l *

Uoper Left Longeron S ta t i on 1 7 5

Osc i l l og raph S t ra in Gage Rea l «

Lower Left Longeron S ta t i on 17 5


Left M a i n Gear B end ing (Fore & Aft)

Osc i l l og raph S t r a i n Gage Rea l »

Fuse lage Structure Lo w er Longeron St a t i o n 65


Left M a i n Gear A x i a l

Osc i l l og raph S t r a i n Gage R e a l »

Left M a i n Gear Shock St ru t

Osc i l l og raph S t ra in Gage R e a l »

Engine Mount Lo w er A ft


Engine

Mount Upper j A ft


[Engine Mount Lower Forward

(Engine Mount Upper Forward

Osc i l l og raph

Osc i l l og raph

St r a i n Gage

S t ra in Gage

* S ee Note at E nd of Table .

Rea l »

Rea l »

3 0 :

.

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TABLE XVII. - Co n t i n u e d

»ARAMETER READOUT TRANSDUCER

St ra in Gage Acce le romete r

SENSITIVITY CHECK

Real« G. Ver t ica l Acce le ra t ion

Osci l lograph

»ilot 's Com- par tment Ver t i ca l Acce le ra t ion

Osci l lograph St ra in Gage Acce le romete r

Real»

' u s e läge Lateral Acce le ra t ion S t a t i o n 286

Osc i l log raph St ra in Gage Acce le romete r

Real»

'uselage Ver t i ca l Acce le ra t ion S t a t i o n 180


Real»

^otor Trans- m i ss i o n Ver t i ca l Acce le ra t ion


Real»

\otor Shaft Vert ica l Acce le ra t ion


Real»

'rope Her Shaft Lateral Acce le ra t ion


Real»

Pylon Lateral Acce le ra t ion

Osci l lograph St ra in Gage Acce le romete r Real»

Pylon Longi tud ina l Acce le ra t ion


Peal»

Vert ical Acce le ra t ion Sta t ion 8?

Osci l lograph

End of Table.

St ra in Gage Acce le romete r

Real»

» See Note at

? 0

"""•"«ä»"«" 6"*"-»' ^

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TA B L E X V I I . - Cont inued

PA R A M E T E R R E A D O U T T R A N S D U C E R S E N S I T I V I T Y C H E C K

Ver t i ca l Accele ra t ion Sta t ion 9-5

Osc i l l og raph St r a i n Gage Accelercmeter

Real*

IVert lca l Accele ra t ion Station Ilk

Osc i l l og raph St r a i n Gage Accelercmeter

Rea l »

Prope l l e r B ox Ver t i ca l Accele ra t ion

Osc i l l og raph St r a i n Gage Accelerometer

Rea l »

Pro p e l l e r B ox Lateral Accelera t ion

Osc i l l og raph S t ra in Gage Accelerometer

Rea l »

p l a p Def l ec t i on Photo Pans l Potent iometer F u l l Throw

»Real : C a l i t 1 wi th

ra t ion by b ias ing known e lec t r i ca l

measurement c i rcu i t resis tance . j

305

- tt...:fi#j**«*P

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TA B L E XVI I I .

I T E M

Forces, Moments, Stresses

S U M M A RY O F M E A S U R E M E N T A C C U R A C Y

O V E R A L L A C C U R A C Y ( Inc lud ing Readout)

M a x i m u m R e l a t i v e Absolu te Erro r E r ro r i

_ of M a x i m u m Value Recorded

Rotor Bla de Pi t ch L ink Force

Rotor Blade F l ap w i s e Moment

S ta t i on ^6 59.5 79.2

105.6 12 4 13 2 158.4 184.8 211.2 237.2

Rotor Blade Chordwise

Moment Station 46

131.5 Rotor haft Bending

Upper 0 ° Azimuth Upper 0° Azimuth Lower 0 ° Azimuth Lower 0 ° Azimuth

Rotor haft orque Rotor haft Lift Propeller ub oment Propeller lade lap

Flapwise oment Station 3.2 16.5 23.1 29.7

Propeller haft orque Turbine orquemeter

Pressure Propeller lade Pitch

Link Force

43.5

646 5 90 639 412 5 5 9 393 360 175 2 3 1 446

1 5 3 0 1 9 0 0

2 8 5 0 2 1 7 0 2 0 3 0 3 6 2 0 9 0 7 0

2 98 2 1 8 . 5

92 .0 181 .0 1 6 4 . 5

2 7 . 9 4 80 36 7

6 . 5 3

U n i t s

L b .

In . -Lb . I n , - L b . In . -Lo . In . -Lb . In , -Lb . In . -Lb . In«-Lb. In . -Lb . I n . - L b . In . -Lb .

I n . - L b . In . -Lb «

P S I P S I P S I P S I

In . -Lb . L b.

In . -Lb .

I n . - L b . I In . -Lb . j In . -Lb . i In , -Lb . ' In . -Lb . In . -Lb .

L b.

1 2 . 6

3.8 4 .8 5 .2 2 .9 5 .1 4 .2 3.1 1.6 3.1 1 .9

7 .1 5 .6

6 .8 5 .2 4 .4 7 .9 5 .1 3 .8 2 .4

1 .6 4 .9 3.8 4.2 2 .5 3.9

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I i

Stresses (Cont'd)

T A B L E XVIII. - Continued

Control lotions

Aircraft

M otions

ITEM OVERALL CCURACY (Including eadout)

Maximum Absolute rror

Upper Left Longeron Stress tation 75 Lower Left Longeron

Stress tation 75 Forward tructure,

Lower eam. Left Station 5

Main anding Gear nee Main anding Gear Axial Main anding ear hock

Longitudinal tick Lateral tick

Rudder edal Collective tick Propeller itch Rudder Vane ngle Rotor lade itch ngle Elevator r im

Deflection Flap Deflection Rotor lade Flap Angle Rotor lade Lag ngle

Pitch Attitude Roll Attitude

Sideslip Angle Pitch ate Roll ate Y a w ate Angle of Attack

.665

• Ml

.122

.27 J » 1.427 1.058

.519

.72

1.0 .238

1.H2

.683

2.35

.56 2.32

.997 2.55 2.32

Deg. Deg. Deg.

Deg./Sec Deg./Sec Deg./Sec

Deg.

Relative Error of aximum Value Recorded

»•2.0 ••3.6

••1.8

••^.7

»•1.2 ••5.2 ••2.2

5.6 6.6

• Calibrating heoretical, accuracy ndeterminate

• Relative rror or tems arked * s n f ull ravel

30 B

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ft

TA B L E X V I I I , - C on t inued

I T E M OVE R i L L A C C U R A C Y (Incls ding Readout)

M a x i m u m R e l a t i v e Absolu te Er ro r Error %

of M a x i m u m Va l u e U n i t a

. 085

Recorded

Vibra- CG, Ver t i ca l G 4.0 t ions Accelera t ion

and P i l o t Compartment . 085 G 4.0 Accelera- Ver t i ca l Acce l e ra t i on

t ions Transmis s ion Case Vert ica l Acce l e ra t i on

N e g l i - g ib l e

C yc/S ec N e g l i g i b l e

P rope l l e r Shaft B ea r ing N e g l i - C yc/S ec N e g l i g i b l e Support Vi b r a t i o n g ib l e i

P rope l l e r Shaft N e g l i - C y c / S e c N eg l i g i b l e Vibra t i on - Ver t ica l g ib l e

P rope l l e r Shaf t Vibra - N e g l i - C yc/S ec N e g l i g i b l e t i on - Long i tud ina l g i b l e

P rope l l e r Shaf t N e g l i - C yc/S ec N e g l i g i b l e Vibra t i on - Lateral g ib l e

M i s c e l - Airspeed (Indicated) 2 . 2 ^ 4 K n . 1.2

laneous Pressure A l t i t u d e 11 .2 F t . 0 .5 Clock . 5 Sec. N e g l i g i b l e

0 .8 urbine R P M 5 0 R P M Outs ide A i r Temperature 2 . 2 ^ Deg. (C ) 0 .8 Fuel F l o w 12 .2 Lb. /Hr. 2 .0 Rate of C l i m b 173 Ft. /Min. 5 .7 Gas Generator R P M 15 6

5 5 . 8

R P M

H P

0 .5

3.9 t ombined Turb ine Power Quant i - (Torque X R P M )

t ies Rotor Power (Torque X R P M ) 14 .1 HP 5 . 1 6 P rope l l e r Power 31 .8 HP 2 . 6 2

(Torque X R P M ) True A i r speed ( I . A . S , , 2 . 5 2 Kn. 1 .3 3

Pressure Alt i tude, O AT ) i

Dens i ty Rat i o . 0 0 8 0 .8

309

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un i t s by u s i n g t h u - c i rcu i t sens i t iv i ty.

The di ff e rence between the In i t i a l and f inal s t a t i c read ings of a paramete r leads to an unce r t a in ty as t o the zero read ing for any bl ip . hen a di ff e rence occurred , the dr i f t was d i s t r i b u t e d l i nea r ly t h roughou t each f l igh t b l ip in the d a t a reduc t ion compu te r program. a t a f rom any c i rcu i t s hav ing a dr i f t of 10 percen t o r grea te r were disregarded . I t »cs cons idered that t he most p robab le d r i f t d i s t r i b u t i o n was one-ha l f of the worst case. he dr i f t e r ro r does no t apply to a l t e r n a t i n g readings .

The on ly probab le er ror o r i g i n a t i n g f rom the osc i l l o -

graphs i s the input f requency e ffec t s on the galvanometer. This effec t was corrected in the compute r program.

Pythagorean add i t ion of errors was used where t he pa ram- e te r i s a de te rmina te func t ion o f the quan t i t y measured . Where there is not a di rec t re l a t i on , the more conse rva t ive a lgebra ic add i t ion was used.

The t a b u l a t e d errors are to ta l errors fo r each pa rame te r ; the percen tage f i gu res re l a t e t o the maximum measurements (sec Table XVIII ) .

All per fo rmance pa rame te rs ( to rque of ro tor, t a i l pro-

pel ler, and engine, al l t a chomete rs , a i rspeed , a l t i t ude and ou t s ide a i r t empera tu re t oge ther wi th t he combined values of tu rb ine , r o t o r and prope l l e r powers , t rue a i rspeed , and densi ty a l t i t ude ) have errors o f 5 pe rcen t or less . n add i t ion , 85 pe rcen t of the r emain ing da ta submi t t ed have probab le i ns t rumenta t ion errors less than 5 percent.

SAMPLE CALCULATION

Roto r shaf t to rque measu remen t e rrors f rom ca l ib ra t i on of 2 Augus t 1956.

Real* 121 K ohms

l ea l* C- 5 3 0

Lever Arm « 258 in. + 1/16 in. - 258 in. +.031

Scale - Cha t l l l i an S/N 2 07 Accuracy +1 .3* at U8 lb.

* ca l is the va l ue ln t housands of ohms of the shunt re- s i s t o r used in ca l ib ra t ion and sens i t i v i t y checks,

leal is the s igna l change caused by the shunt re s io to r expressed in degrees of ga lvanomete r def lect ion .

3 1 0

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A P P L I E D NE T AV E R A G E R E A D I N G W E I G H T M O M E N T AV E R A G E G Ä LV O G A LV O N U M B E R (LB) ( I N V L B ) M O M E N T D E F L E C T I O N D E F L E C T I O N

1 i » 3 11 0 9 1 1 0 8 3 6 1 . 7 9 1 1 . 9 1 5 2 6 i | 1 6 5 1 2 U286 2 . 9 7 5 3 . 1 5 0 3 8 i » 2 1 6 ? 2 2 2 9 6 2 1 . 3 5 8 1 . 8 9 3 1 105 2 7 0 9 0 2 8 2 5 1 5 - 7 0 2 6 . 1 9 7 5 119 3 0 7 0 2 3 11 7 6 6 . 7 0 5 7 . 0 3 6 6 135 3^ 830 3 5 6 0 1 7 . 5 6 3 7 . 9 7 8 7 1*8 38181 3 8 1 8 1 8 .1 3 7 8 . 137 8 141 3 6 3 7 8 8 . 2 9 2 9 125 3 2 2 5 0 7 .3 6 7

1 0 111 2 9 11 2 6 .6 9 2 l i 9 1 2 1 2 5 2 5 . 1 2 7

1 2 7 0 1 8 0 6 0 3-925 1 3 1 1 1 0 5 7 8 2 . 0 9 5

These readings are averages of Increas ing and decreas ing readings.

Curve f i t t ing by method of averages .

Total of f irst four average moments 1 0 8 3 6 1 7 2 8 6 2 2 9 6 2 2 8 2 5 1 7 9 3 5 5

Total of f i rs t four average def lec t ions 1 .9 1 5 3 . ^ 5 0 1 .8 9 3 6 . 1 9 7

Using equat ion of s t raight l i n e

Moment na X def l ec t i on + b

Where n i s the number of readings 1

7 9 , 3 3 5 ■ l a X 1 6 . 1 8 5 + b

Total o ' * second four average moments (us ing number 1 twice because of odd number of readings) 2 8 2 5 1

31176 3 5 6 0 1 38181

1 3 3 5 1 5

3 1 1

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Tota l o f second f o u r ave rage de / lec t lons 6 .197 7 .035

8 . ' ? 3 7

133 .515 - * < & X 2 9 . 6 4 8 + b

S o lu t ion o f the two s imul taneous equa t ions y ie lds

Moment ■ Mll6 Deflec t ion + 2870

In t he I c a i value 20 .53° , t he moment ■ 84 ,501 4.n.-lb.

Value of augment used - 81 ,900 - lb.

Erro r in s lope - . ,501 -JjUaJM • 3 .08J

Assume 3- inch t race , f o r two read ings w i th in .01 inch

Readout error - 2 ^ ^ 1 - 66«

Overal l accuracy f o r ca l ib ra t ion

.03 l eve r a r r * . l eng th er ror 1.30 sca le e r ro r 3 .08 s lope e r ro r

.66 readou t e r ro r 5 . 0 7 percent

No res i s to r change between ca l ib ra t ion and f l igh t s , the re fo re no er ror f rom th i s source .

Er r o r i n s ta t i c read ing . 0 1 in. Erro r i n t race read ing . 02 in.

Error in leal read ing +.01 in.

Probab le er ror in read ing i.Ol) 2 + ( .02) 2 + ( .01) 2

- . 0245 in.

Drif t ■ .0085 in. ( f rom random samples) S ens i t i v i ty - 43 .323 In.- ih/ln. def lec t ion

Maximum t o rque » 1 9 2 , 0 0 0 in.-ib, ( reference Figure 149)

Reading error 0 2 i ?A% ^3 2 ^ -553

Drif t er ror

Overal l er ror

. " 3 , 3 2 3 192,000

.0085 X 4 3 . 3 2 0 - .192* 1 9 2 , 0 0 0

y(5.07) 2 .553) z .192) 2

- .11%

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TABLE X X . ESTIMATED D R A G BREAKDOWN, 1 6 H - 1 A

A P P L I C A B L E QUIVALENT I T E M PARASITE N E T W E T T E D

A re a ( F i ' 2 ) Area (F T2 ) Fuselage 6 . 2 7

A. Drag

B a s i c Fuselage 2 . 5 0 408 Surface Imperfections . 1 0 Blisters, G a p s . 0 6 N o s e Boom . 2 1 N o s e Weights . 2 9 Antenna . 0 4

P y l o n . 6 6 4 2 Engine I n l e t . 2 8 Engine E x h a u s t . 6 0

B. Momentum L o s s e s j O i l Cooling F an . H O

Engine N et D r a g • 35 , Fuselage Ventilation . 7 8

Landing Gear . 9 5

Main Gear ( Op e n We l l s ) . 6 1 Ta i l Gear . 3 4

Rotor 3 .15 H i n g e Assembly 2 . 8 5 B l a d e S h a n k s . 3 0

Wi n g 1 .2 0 140

R i n g Ta l l 1 .9 b Duct ( I n c l u d i n g S u r f a c e . 9 5 9 6

Imperfection) R u d d e r s ( F i t t i n g s , . 6 0 4 9

B o l t s , etc. ) Elevators and Horizontal • 1 7 1 8

Stabilizer S t a t o r s . 2 1

1 3 . 5 3

33

TOTALS 786

i Ö T E : W h e e l s and F l a p s U p. f c

-

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• ;

APPENDIX I I I

STABILITY AK D CONTROL ANALYSES

LONGITUDINAL CONTROL POWER A N D ELEVATOR TRIM POWER IN HOVER

S t i ck pos i t i on requ i red fo r t r i m has been measured in the hover co nd i t ion for var ious cen te r-o f -g rav l ty loca t ions and e leva to r def lec t ions . hese resu l t s are shown in Figures 179 and 180.

The fa i red curves of s t i ck p o s i t i o n versus cen te r-of - grav i ty loca t ions and versus e l e v a t o r se t t ing are cons is ten t with t he f ina l level f l igh t s t i ck p o s i t i o n versus speed curve which i l l u s t r a t e s t he l 6H - iA long i tud ina l s t a t i c s t ab i l i t y (Figure S * « ) .

The ra te of change of s t ick grav i ty sh i f t shown in t he plo t ieri raental iy conf i rmed es t ima te d i s c u s s e d in the hover response essen t i a l ly independent o f gross fo l lows:

d isp lacement with c e n t e r- o f - i s subs tan t i a tec by t he ex- of l ong i tud ina l con t ro l power ana lys i s . his parameter,

weight j is computed as

M« - ' 1 , 500 f t - lb / inch , based on W « 6000 lb

dx _1 12 M ,

(73)

-6 .000 12 x 1 ,500

■ - 0 .3 3 3 inch s t i ck p er inch cen te r-of -grav l ty d is - placement

With th is parameter, the e leva to r t r im effec t iveness in hover cam be de te rmined in te rms of both cen te r-of -grav l ty d isp lacement and p i t c h i n g moment . These parameters are found as fo l lows :

dx d6 e

where

di d6

d6/d6,

d6/dx

-0 . 0375 ln/deg. f rom Figure 180

(7^)

e

3 2 2

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-

CG. P O S I T I O N - I N C H E S F O RWA R D T T T

O F S H A F T F igu re 179« Long i tud ina l St i ck P o s i t i o n versus

C . G . P os i t i on - Hover ing , Neut ra l Eleva to r,

3 2 3

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KIMU MJ. :e*£*im>t**

NOSE ? « O SE O W N E L E VAT O R T R I K - D E G R E E S

F igu r e 180. Longi tud ina l S t i c k Po s i t i o n versus El ev a t o r Tr i m - H o v er i n g Fl igh t .

3 2 4

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dx

die -0.0375 -0.333

0.113 nch enter-of-gravlty isplacement per degree f levator.

Thus the levator s capable f alancing .'J-inch center-of~gravity hift in over t ixed tick osition. The corresponding itchlng-moment effectiveness is

V = -M6 fr (75)

-(-l J 500) ( - . 0375)

-56.3 t-lb/deg

A N G U L A R A C C E L E R AT I O N I N H O V E R

Control power, expressed i n terms of i n i t i a l angu lar ac - ce le ra t ion per inch of control mot ion , has been ev a lu a ted from hover control response tests . An g u la r rate traces were used to obta in control power considered as th e ra t io of control moment pe r inch to the i ne r t i a about th e cor responding motion axis .

T he resul tant control power i s g iv en by the slopes of the plo t ted angular ve loc i t y data (Figures 97, l 8 l and 3 8^). T he exper imenta l values for these tests are compared wi th th e exper imenta l resul ts from th e 50-knot dynamic tests i n the f o l l o w i n g table, together wi th th e minirrum requirements based on M I L - H - 8 5 0 1 A hover sp ec i f i ca t io n (assuming a pure i n e r t i a system), and i s summarized below.

C O N T R O L P O W E R ( D E G / S E C2 P E R IN . )

H O V E R

5 0 - K N O T D Y N A M I C

T E S T S

M I N I M U M H O V E R R E Q U I R E M E N T, M I L - H - B 5 0 1 A

P i t ch -9.2 -7 .8 - ' 1 . 7

R o l l 5 1 . 0 5 3 .8 11 .2

Y aw (Right Pedal) -3 0 .0 -33.^ -11 . 5

Y aw (Left Pedal) -ih.H -23 . 9 -11 . 5

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♦ 1

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I t i s seen t h a t the comparison of t h e test-program r e - s u l t s with t h e MIL-H-8501A specification I s quite satis- f a c t o r y . t I s evident t h a t t h e control power about a l l a x e s satisfies t h e minimum requirements.

ANALYSIS O F D Y N A MIC S AND C O N T R O L R E S P O N S E I N HOVEq

P i t c h Response

A theoretical analysis o f t h e 16H-1A control response characteristics I n hover h a s been m a d e . Calculated responses f o r t h e

t e s t control

I n p u t conditions

have

been

compared

t o

t h e measured r e s p o n s e s . ethods o f solution o f a l l response equations a r e taken from Reference 1 , a n d the longitudinal equations of motion a r e given I n Reference 7 . O n t h e basis of t h e c l o s e agreement obtained, t h e estimated values f o r control p o w e r , damping, velocity stability, and the resultant stability r o o t s were verified. I t I s shown t h a t the 16H-1A design exhibits a typical helicopter c u b i c characteristic equation with a short-period t i m e constant and phugoid characteristics comparable t o those o f single-rotor heli- c o p t e r s b u t with a longer period. ontrol power exceeds the minimum MIL-H-85OIA requirements.

T h e hover equations o f motion i n dimensional f o r m , e x - cluding t h e uncoupled plunging m o d e , a r e given a s follows f o r t h e pitch angle a n d velocity degrees o f f r e e d o m .

( Xv -mD) V - ( m g ) 6 » - Xö 6 ( t )

7 6 )

( Mv ) V + ( Mq D - I y D2 ) 8 - - M6 ö ( t )

7 7 )

Solutions of these equations for t h e pitch angle response d u e t o stick i n p u t , with z e r o initial conditions, yields the following third-order response with a lead time constant t e r m .

K(D + i - ) 6 { t ) 6 - * •

7 8 ' ) ( D + =i-)(D2 + 2C t oD + « 2 )

l 0 i

w h e r e K = ü i . I y

T ^ s p

T h e i n v e r s e lead t i m e constant i s

1 x 6 ^v x v

fTrt ,

l Hr x M^- T

79)

339

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1

and he haracteristic quation s iven y

(D i—KD 2 + cU o D .02 ) » 3 - ( }a.) D2 +

ls p

g M

^ j) D (-17) Xv M q

(80)

Estimates f he tability nd ontrol erivatives fcr

hover re

Xv 5.3 b-sec/ft X6 80 lb/in Mv 1» .1 ft-lb-sec/ft M q 3,8AO ft-lb-sec/rad M ä 1,500 ft-lb/ln ly 1,000 slug-ft 2

m 86 s lugs

T he r e su l t i ng l ong i t ud ina l dynamic parameteio are

K -7 .8 0 deg/sec 2 per inch

i- » 0 . 0 0 6 1 se c - 1 TL

_L. 0.622 ec" 1

•Lsp

c = 0.305

< « i0 = .403 ad/sec

and he oscillatory ode eriod s found s follows

2*

.-.21. 0.103 \/l-(0.305) ;

(81) T/

= 16.3 ec.

31 0

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I

T he long i tud ina l control Inputs and corresponding pi tch angle responses for Tes t 3 3 5 , t J l i p s 3 and k (F igures 65 and 66) , have been u t i l i z ed for the cor re la t ion of the dynamic characteris t ics . I n these tests, the control motions are best represented by success ive pulss Inputs. Im p u l se responses of v a ry in g magnitudes and I n i t i a t i o n t imes are therefore adopted I n th e ana 1 , 'sls. T he bas ic response equat ion for a unit impulse input of 1 inch-second i s g i v e n i n degree units , i n terms of th e Lap lace var iab le , as

7 . ^ - 0 ( s + . 0 0 6 ^ )

e(s) = ( s+0 . 622) [ s 2 + 2( -0 . 305X0 . '<03)s + (0 .H 0 3 ) 2 ]

The nverse ransform, or ime olution, is

6 = .80 -0 - 6 2 t - -75 0 "1 23 t :Ln(0.385 + .77)

Tes t 3 3 5 , 3 11 ? 3 (F igure 6 5 ) i s adequately represented by i n d iv id u a l Jrpuis of 1 .8 inch-second at t 0 . -0 .5 inch- second a t t ■ 1 c^cond, and -2 .0 inch-seconds at = 2 seconds. T he total response produced by the three inputs w as calculated us ing th e above so lu t ion by super impos ing the i n d iv id u a l responses weighted by tne i r respec t ive Im p u l se magnitudes. he same procedure was employed for Test 335, B l i p h (F igure 66) , wi th impulses of -1 .1 inch-seconds at t * 0 , O . l J inch-second at t » 1 .5 seconds, and 1 .3 inch- seconds at t 2 .5 seconds. hese responses are p lo t ted together wi th th e test measurements, and good agreement i s shown. hese assumed impulse magnitudes are sm a l l e r than the in tegrated pulse areas shown i n tn e actual s t ick mot ion curves i n order to correct for the effects of control sys tem e l as t i c i t y present on only the early f l i gh t s (see Technica l Problems sect ion) .

Rol l Response

Rol l responses i n hover are g iv en by the data of Tes t 3 3 5 , B l i p s 5 and 6 , (Figure 97). F rom these tests and the lateral dynamic analys i s at 5 0 knots, a l a tera l control power to i ne r t i a rat io of approx imate ly 50 degrees per second pe r second per inch has been es tab l i shed . I t i s est imated, based on the 50-knot lateral dynamic response resul ts , that the steady-state rol l ve loc i t y per i nch i s approx imate ly 39.0 degrees per second i n hover. T h i s l eve l of i n i t i a l rol l accelerat ion an d maximum ro l l v e lo c i ty i s f a i r l y ev iden t i n the h o v e r responses where only r e l a t iv e ly small s t ick def lec t ions from t r im are required to produce momentary ro l l rates of 1 0 to 15 degrees per second.

3^1

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I t should be noted that strong la teral control was pro- v ided I n th e 1 6 H - 1 A i n order to ^nsure adequate cont ro la- b i l i t y n the unloaded rotor high-speed f l igh t condi t ion. U se of properly proport ioned a i l e ron control i n conjunct ion wi th cy c l i c can be u t i l i z e d to obta in constant lateral control power characteris t ics .

Y aw Response

Y aw responses i n hover have been analyzed on th e bas i s of a s ing le-degree-of - f reedom sys tem us..ng es t imated s t ab i l i t y and control der iva t ives . he measureu responses are shown i n Tes t 335, B l i p s 7 and 8 , (Figures 1?4 and 195 ) , together

wi th the t heo re t i ca l l y der ived yaw responses which ar e seen to ver i fy the es t imated der ivat ives . rom th i s corre la t ion , the d i rec t ional character i s t i cs i n h o v e r ind icate an i n i t i a l y aw accelera t ion to the r ight of 30 degrees per second per second per inch, w i t h a steady-state yaw rate of 1 88 degreec per second per inch. or y aw to the left , th e i n i t i a l accelera t ion I s Ib.U degrees per second per second per inch wi th a steady-state y aw rate of 9 0 .5 degrees per second per Inch. he response t ime constant i s 6 .3 seconds, and the I n i t i a l accelera t ions i n both di rect ions exceed th e m i n i m u m M I L - H - 8 5 0 1 A requirements .

T he equat ion of mot ion i n yaw i s

( I Z D - Nr ) D * = N6 6(t)

S o l v i n g for th e yaw rate response g i v es

(82)

D K 6(t)

T (83)

where

and th e i nverse t ime constant i s

1 T

- N ,

Est imates of th e d i rec t ional response parameters are

Nr = - 1 , 6 4 0 f t - lb-sec/ rad Nx = - 5 , ^ 0 0 f t - lb / inch (r ight pedal)

- - 2 , 6 0 0 f t - lb / inch (left pedal) I z « 1 0 , 3 0 0 s lug-f t 2

3^2

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Figure 194. Y aw Response i n H o v e r - Lef t Yaw, C.G . 8.5,

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' J 8 10 11 12 T I M E - S E C O N D S

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3 1 ) 4

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T he overa l l stat ic s t ab i l i t y, proport ional to tn e s tat ic margin , can be expressed i n terms of the s t ick posi t ion wi th a i rspeed s lope:

d6 dV

1 -M6

[MV + C L v ( n0 - XC G ) ] (85)

Est imates of th e above parameters for steady l ev e l f l igh t operat ion at 15 0 knots have been made and ar e g i v e n as fo l lows

M6 * -1 3 0 0 f t - lb / inch Mv « 3 7 f t- lb /f t /sec

C = 5 feet L y = ^1 lb/ f t /sec no 0 . 5 1

AC G - 0 - 2 7

T he s t ick posi t ion s lope i s then

d« W [3 7 + 5 x 2 1 ( 0 . 5 4 - 0 . 2 7 ) ] (86) 30 0

» .050 in/ft/sec

» 0 . 0 8 5 in /knot

This va lue i s seen to be wel l represented by the slope at 15 0 kno ts of the fai red curve for s t ick pos i t ions , thus ver i fy ing the es t imated der iva t i ves . he pos i t i ve s lope i nd ica t ing p o s i t iv e overa l l s ta t i c s t ab i l i t y i s ev iden t th roughou t the operat ing reg ime above about a 70-knot forward speed.

T h i s ana lys i s was for test data obtained at forward cen te r of grav i ty (9.6 inches) . he long i tud ina l s ta t i c s t ab i l i t y for the af t center-of-gravity pos i t i on (3-1 inches) can be eva lua ted by th e v e r i f i ed approach g i v e n above. he only parameters which are center-of-gravity dependent are M y and X Q. At the aft pos i t ion , M y = 16 foot-pounds per foot per second and XC Q ■ 0.37 .

Thus,

d ö dV [1 6 + 5 x 21 ( 0 . 51 - 0 .3 7 ) ] 1300

0 . 0 1 9 in /f t /sec

0 . 0 8 3 in/knot

31 6

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>

T he reduct ion i n s t a b i l i t y due to center-of-gravlty shi f t i s compensated by the increase i n ve loc i ty s tab i l i ty. I t should be r ea l i zed that t h i s tendency for af t center-of- gravi ty sh i f t to have only m in o r effect on th e l ong i t ud ina l control pos i t ion gradien t i s typ ical of hel icop ter character- i s t i cs (Reference 2, pages B1*2* and ^ 58 ) . ;t these speeds he 1 6 H - 1 A un load ing i s such that t h - » rotor i n f l u - ences govern this s ta t i c s t ab i l i t y parameter rather than the pure f i xed -wing cont r ibut ions .

r '

Direc t i ona l

T he steady s i d e s l i p pedal pos i t i on curves of F igu re 10 3 i l l u s t r a t e d i rec t l y the p o s i t iv e direc t iona l s ta t i c s t ab i l i t y of th e 1 6 H - 1 A at al l speeds above hover. I n addi t ion , these pos i t i on curves a l l o w the computat ion of the stat ic s t ab i l i t y der iva t i ve Na u t i l i z i n g es t imated direc t iona l control power values. Control power at h o v e r has been v e r i f i e d by the control response ana lys i s and cor re la t ion g i v e n prev ious ly, and control power es t imates at 5 0 knots ar e v e r i f i e d i n the forward speed dynamic ana lys i s i n a f o l l o w i n g sect ion of t h i s appendix .

T he steady s i d e s l i p e q u i l i b r i u m equation i s

Nß x 0 + N x 6 = , ( 8 7 ) g i v i n g

6 -H-^ F or s ides l i p at 50 knots, the es t imated control power

der ivat ives are

N6 = - 6 , 0 0 0 f t - lb / in . (r ight pedal)

= -i | ,300 f t - lb / in . ( le f t pedal)

and rom he osition urve, he lopes are een o e

H =0.92 in / rad ( r ight pedal)

=1.26 in / rad ( left pedal)

Apply ing these measured slopes y i e l d s

Nß = - 0 .9 2 ( -6 ,0 0 0 )

5 , 5 0 0 f t- lb /rad

3^7

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and Nß 1.26 (-U,300)

« ,^00 ft-lb/rad

with n verage alue f g J1*50 t-lb/rad

At 10 0 knots, the es t imated control power i s symmetr ical w i th a Va lue N r- - 9»800 foot-pounds per inch. he pos i t ion curve s lope at th is speed i s

HI » 2 . 5 2 i n / r a d UP

giving irectional tability alue f

Nß 2.52 (-9,800)

= ^,600 ft-lb/rad

D Y N A M I C S TA B I L I T Y A T 5 0 K N O T S

Longi tud ina l

A theoret ical ana lys is has been undertaken i n order to es t ab l i sh the major parameters govern ing th e l ong i tud ina l h an d l i n g q u a l i t i e s of the 1 6 H -1 A . hese resul ts have been ap p l i ed to the t rans ient responses obtained i n the test program. A su ff i c i en t ly h igh degree of correlat ion i s shown to ve r i f y ehe ana lys is procedures and s t ab i l i t y estimates. I t i s demonstrated that at a l ow forward speed of 50 knots (the hover regime considered to apply up to 30 knots) , the 1 6 H - 1 A design e xh ib i t s a w e l l damped shor t -per iod mode w i th a per iod of ^ .8 seconds and a c r i t ica l damping ra t io of 0 . 5 7 . he s t i ck - f i x ed neutral poin t i s at 5 1 percent of th e mean aerrdynamic chord. he control power to i nerc la r a t io produces an i n i t i a l p i tch accelerat ion of 7 .8 degrees per second per second per inch an d the steady state pi tch veloc i ty i s ^ .7 degrees per second per inch. T he p i tch rate response meets th e i n f l ec t ion point t ime requirements of M I L - H - 8 5 0 1 A .

T he s t ab i l i t y - ax es equations of mot ion i n dimensional form ar e g i v en as fo l lows for th e shor t -per iod control response, i n w h i ch the f l igh t speed may oe considered constant

[L a + (mV0 ) D] a - ( Fn V0 ) 0 6 -L 6 6(t)

(Ma ) a + (Mq D-I y D2 ) 6 = -M6 6(t)

3^ 8

(88)

(89)

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*

1

Solut ion of these eauations for the pitch angle response due to stick Input, with zero In i t ia l condit ions, yields the fol lowing second-order rate response with a lead t ime- constant t e r m .

K ü + i ) 6(t) e = D ( D2 + 2 ; a )0 D + u .0

2 ) ( 9 0 )

where

= « .

9 1 )

I y (Inverse Lead i _ L

a Mq L 6

.

Time Constant) T S mV0 " M6 mV0

92 ^

Iy 7jr- MQ

(Critical r „ a n C /

—

93) Damping atio) L , M M u a

2\/l y (-Ma-M q ^ )

fMa Mc ' (Undamped

"o-"q ^j-

Natural

o /

9^)

Frequency)

y

Estimates of he stability erivatives for teady level flight at 50 nots are

La 7,000 Ib/rad L6 ^20 lb/inch

Ma 20,000 ft-lb/rad

Mq 6,800 ft-lb-sec/rad

M6 1,500 ft-lb/inch I 1,000 slug-ft 2

mVj 5,000 lb-sec

The resulting esponse parameters are

K = 7.80 deg/sec 2 per nch

sr 1.513 ec -

C = .570 ü )0 1.59 rad/sec

3^9

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(

H

o

S3H0NI

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dn 3 S 0 N NQ aeon oas/oaa

- 3 . LV H H D i l d

d f l 3 S 0 N 'N O 3 S 0 N S 3 3 H 0 3 a

^33DMV H O i l d

QM S 3 H 0 N I - H O X I d

N O I J J S O d M O I i S

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S 3 H O N I - T - a i O H ' N O I U S O d M O I i S

d f l 3S0N ' N O 3S0N 032/030

- 3 J , V H H O i l d

d f l 3 S 0 N ' N Q 3S0N S33H03a

- BI O N V HOJJd

ßM S 3 H O N I - H O . L I d

N O I J i l S O d M O I i S

351

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■

T he long i tud ina l control input and pi tch angle response f or Test 3 1 5 , B l i p 5 , (F igure 69) , have been u t i l i z e d for the corre la t ion of the dynamic character i s t i cs . ere, for convenience , th e control i npu t i s rapresented by a s ing l e i m- pulse- type input of 0 . 7 5 inch-second magni tude I n i t i a t e d at zero t ime. he pulse durat ion i s short enough compared to the system per iod to a l l o w th is approach. ak ing th e L a- place t ransform of the p i t ch response, i n degree uni ts ,

e(s) -.75 .8 0 (s > • . 513}

s[s 2 + (0.5"7)(1.59) s+(l.591*1 (95)

The inverse ransform, or ime olution, is

-0.9l tft = 3-50 .056 sin(l.31 + .1) (96)

T h i s respons, s plo t ted w i t h the test measurement and shows very c l o s * » greement for the f i rs t 3-5 seconds of the t r ans ien t . After this t ime, the r e l a t i ve ly smal l correct ive s t i ck mot ion evident i n the control trace causes the actual response to d ive rge toward the i n i t i a l pi tch att i tude.

Latera l

A theore t ica l analys i s , s i m i l a r to that of the longi tudinal case, has been made for the pure rol l response to la t- eral s t ick motion. I n these runs, th e s t i ck was pulsed l a t e ra l l y i n alternate di rect ions so as to produce ro l l ex - cursions which were rather quickly returned to tr im. A s a resul t , the usual at tendant yaw and s i des l i p mot ions were small enough compared to the ro l l response to be cons idered n eg l ig ib l e i n th e ana lys i s .

T he s ing le-degree-of - f reedom ro l l equat ion of mot ion i s

( I X D2 - Lp D) * = L6 6(t)

97)

S o l v i n g th is equation for the ro l l angle g ives K6(t)

D ( D + ^ ) (98)

where LAK = ^

and T, th e f i rs t order response t ime constant , i s g iven by

1 Lx

352

iiummat HmiiiM

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i

T he est imates of the la teral parameters at a f l igh t speed of 5 0 knots ar e

I x = 1 , 6 0 0 slug-ft 2

Lp = - 2 , 5 3 0 f t - lb-sec/ rad L g = 1 , 5 0 0 f t - lb / lnch

resu l t ing i n a la teral control ga in and inverse t ime constant of

K = 53 . 8 deg/sec 2 per inch

fv 1 . 5 8 sec 1

Consider ing a unit im p u l se input of 1 inch-second, the Laplace var i ab l e response equat ion i s

♦ (s) = ^3 .8 s ( s+1 .5 8 ) ( 9 9 )

and i ts inverse i s

4

4 3^.0 ( l -e- 1 - 5 8 ^ ) ( 100 )

T wo lateral test responses were s tud ied i n the analys i s . I n Tes t 315 , B l i p s 7 and 8 , th e former i s adequately represented by a -0 .26 - inch- second impulse at zero t im e and a correct ive impulse of 0.M inch-second taking p lace at 1 .2 5 seconds; the la t ter input I s a 0 . 2 1 inch-second impulse at zero t ime fo l lo wed by a cor rec t ive im p u l se of - 0 . 2 8 inch- second appl ied at 1 .2 5 seconds. These assumed im p u l se magni - tudes ar e sm a l l e r than the in tegrated pulse areas shown i n th e actual s t ick motion curves i n order to correct for th e effects of control sys tem e la s t i c i ty present on only the ear ly f l igh t s (see Techn ica l Problems sect ion). T he total response due to both inputs was calcula ted for each case using the above so lu t ion and super impos ing th e i n d iv id u a l responses weigh ted by the i r respec t ive impulse magnitudes.

T he r esu l t ing equations for both cases are:

Test 315 , B l i p 7 (F igure 198) :

From t = 0 to t = 1 .2 5 seconds

* = -8 .84 (l-e ' 1 ' 5 8 * ) (101)

353

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t dfl S0N 'N Q SSON

C

S3F10N I

-HOild 'NCIilSOJ MOIXS

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'MOIilSOd M O I i S

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-3TJNV i J O i l d c

S 3 H 0 N I

- H O i l d ' N O I I I S O d XOIi IS

0 3 S / D 3 a S 3 3 y ü 3 a a —- S 3 H 0 N I - * • i

-33oa ' N O I l I S O d M O I i S

3 5 5

. ,c*wii Wi*aaÄ«w«5w«^**- ' -

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a n a s o N Nd 3S0N S33H03a

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03S/fj3a - 3 J , v u lion

S33H03a CJ3 I0NI-^ 1 - I TO H

' N O I i l S O d MOUS

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a n a s o M Na 3SON " • " " " S33H03a ~ * ' 'G M 3H0NI -HOIId 'NOIilSOd MOIXS

r - » m

032/030 -3 iVH l l O H

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S 3 H 0 MI -non

N O I X I S O d M O I i S

357

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I i

Fro m t 1.25 seconds, continuous

4 - 6 .12 ( 1 - 16 .1 e -1 - 5 ^

Test 315, Bl ip 8 (Figure 199) :

From t " 0 to t « 1 , 2 5 seconds,

* » 7 . I f (l-e- l - 5 8 t )

From t 1 . 2 5 seconds, continuous

« .38 (1-25.8 - - 5 8 *)

(102)

(103)

(1014)

These responses are included n he lots of he easured roll esponses. It can e een hat correlation s atis- factory for he first 3-5 econds in lip (Figure 192) and or more han econds in lip (Figure 199)-

The bservations rawn rom he verified tability nd control haracteristics are ased n the usual onsideration of -inch tep nput. The teady-state oll ate er inch f tick otion s 3^.0 egrees er econd. This results in helix ngle er nch f & . .070. In ddition,

2V the s teady s t a t e is reached, essen t i a l ly, iu 3 t ime constants

or

3T = Y^ «1.9 seconds,

w hic h i s a s u f f i c i en t l y r ap id response t ime. Direc t iona l

Direc t iona l dynamics have been i nves t iga ted i n analyses of f l igh t t es t records of th e d i rec t iona l response to input con t ro l mot ions . In the 50-kno t response data,the aircraf t was s t ab i l i zed in pi t ch and ro l l so that only d i rec t iona l motions i n which t he s i d e s l i p angle i s equal and opposi te to yaw angle resulted. I t shou ld also be pointed ou t that coupled l a t e ra l -d i rec t iona l dynamics are mi n i ma l at 5 0 knots,

du e es s en t i a l l y to the favorably l ow l e ve l of ne t dihedral effert of the 16H - 1A configurat ion.

Analyses have been performed by first dete rmin ing the d i rec t iona l s t ab i l i t y and control der iva t ives and then ap- p l y i n g these to th e aircraf t equat ion of motion. he p r inc i - pa l factors, control power and sta t ic s tab i l i ty, have been es tab l i shed by the data correlat ions g i v en prev ious ly i n th is appendix.

358

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The single egree f reedom esponse quation n aw s

(I Z D2 - ir ß ) * Jä 6(t)

105)

and ne resulting xpression for he aw ate esponse

to edal displacement time function is

^

* _*e- üt i

1 o6)

D + 2;CJ 0 D D0Z

where

.'

_

or,

z

C =

-u I

LZ

$ = 108) V^ ß

(109) T,

The tability nd ontrol erivatives, as previously obtained, are

N6 = 6,000 ft-lb/in (right pedal)

= ^,300 ft-lb/ln (left edal)

I z 0,300 lug-ft 2

Nß ,^50 t-lb/rad

and he stimated irectional am; ng s

Nr -A, 600 ft-lb-sec/rad

With tnese arameters,

K 33. deg/ssc 2 per nch (ri;ht pedal)

= 23-9 deg/sec 2 per nch (left pedal)

r . = .307 u )0 = .726 rad/sec

Also, the damped atural frequency is

u ) = JJ/TT2

110)

= 0 .7 2 6 v/ l - ( 0 . 3 0 7 ) 2

= 0 .6 9 0 rad/sec

35 9

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1 and t he corresponding osc i l l a t o ry per iod I s

P - £1 u

(111)

- 6 . 2 8 0 . 6 9 0

■ 9 .10 seconds

These results Indicate a moderately damped long-per iod o s c i l l a t i o n w i t h h i g h control effect iveness I n yaw at l ew speed. t moderate and h igh speed, the response I s of shorter per iod.

T he control responses to be s tudied are the 50-kno t pedal Inpu t cases for P l l g n t ^ 3 5 , B l i p s 1 and 2, shown I n Figures 2 0 2 and 203 . I n these p lots I t i s seen that th e control mot ions are pu l se type. ecause of the long o s c i l - la tory per iod compared to the actual pu l se durat ions , an Impulse approach to the theore t ica l response I s app l i cab le . n ad - dit ion, lengthy response corre la t ions up to 1 5 seconds are feas ib le due to the previous ly ment ioned condi t ion of pure y aw response.

T he yaw rate response. I n degrees-per-second uni ts , for a uni t Impulse funct ion (1 Inch-second r ight and le f t pedal input) i s g i v e n i n the Laplace v a r i ab l e as

♦ (s) K s s 2 + 2 ( 0 . 3 0 7 ) ( 0 . 7 2 6 ) s + ( 0 . 726 ) 2

and th e resu l t ing response t ime so lu t ions are

* « 35 .1 e- 0 - 2 2 3 t s i n ( 0 . 6 9 0 t + 1 . 8 8 )

for r ight pedal uni t impulse disp lacement from t r im and

I = -2 5 .1 e- 0 - 2 2 3 t si n ( 0 . 6 9 0 t + 1 .88)

for lef t pedal uni t impulse inputs.

(112)

(113)

(111)

Ex am in a t io n of the actual control pulses Ind ica tes that a representat ive input for Tes t 4 3 5 , B l i p 2 (F igure 203 ) , i s -0 .5 Inch-second and for Tes t I 35 B l i p 1 (F igure 2 0 2 ) , i s 0 .7 inch-second. U se of the above so lu t ions , weigh t ed by t he i r respect ive incut magnitudes, y i e l d s the theoret ical responses shown i n Figures 2 0 3 and 2 0 2 . he expected d i f - ferences between th e i n i t i a l spikes because of the f i n i t e

3 6 0

i'wiwingiinMBft Mwktm*»**um*~

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H O J J d N O L L I S O d M O I i S •Q MJ- ^ 3 H 0 N I

7 1 0 U N O I i l S O d X O I 1 S S 3 H O N I - T

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a - « - o a s / o a a

- 3 I V U . M VA

- » - a - 10iMV Oild

H O 3 1 O N V a a o a

- S 3 H 0 N I - K O I I I S O d

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I , , , .

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CI M d - « — S 3 H 0 N I - H O J - I d

' N O I i l S O d M 3 I J . S

SHHONI—- r

H O I i l S O d M O I i S

0 3 5 / 0 3 0 - 3 i LV U M V A

d f l 3 S 0 H -N d H S O M 1 U O H H O T •DM - 3 1 0 N V H O J - I d

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7 v a3 d HsaanH

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362

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j

rate of control applicat ion required in any actual system are apparent i n the comparisons between the actual and theoretical responses. Nevertheless, t h e response correlations are satisfactory and they substantiate th e der iva t ive evaluations a s well a s the corresponding direct ional dynamic characteristics described.

A N A LY S E S OF R E S U LT S A T 15 0 KNOT S

Longitudinal Dynamics

• Control responses obtained a t 150 knots have been analyzed In a s imi la r manner t o that used for the 50-knot dynamic tests presented i n this appendix. A g a i n , a suf- ficiently close agreement i s shown so that the analysis approach and stabi l i ty der ivat ive estimates are considered verified. I t i s shown that the 16H-1A a t 1 5 0 knots ex - hibi ts a damped short-period mode with a period of 2 .0 seconds and a critical damping ratio of 0.^3« T he stick- fixed neutral point i s a t 5 percent of the mean aerodynamic chord. he control power to iner t ia ratio produces an ini t ia l pitch accelerat ion of 6.8 degrees per second per second pe r inch and the steady-state pitch veloci ty i s 2 .^ degrees per second pe r inch. T he normal accelerat ion and pitch

rate response meet th e inf lect ion point requirements of MIL-H-8501A.

The stabil i ty-axes equations of motion i n dimensional form are g iven a s fol lows for th e short-period mode response, i n which the fl ight speed may be considered constant.

[ L0 + (mV0 ) D] a -(mVo)De = -L 6 6(t) (115) 1

( Ma ) a + ( Mq D -I y D2 ) e = -M6 ö(t) ( l l6)

Solut ion of these equations for the pitch angle response due to stick input, with zero In i t ia l conditions, yields the fol lowing second-order rate response with a lead t ime constant term.

4

(D+i) 6 ( t )

9 = D(D2 + iT^D + a .02 )

117)

where

= T-

118)

Inverse ead

La Ma 6

Time Constant

vo 17^

y

363

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C r i t i c a l Damping Ratio

Ly v; - Mr

/"= ^ / IyC-Ma-Mq. ^

(120)

Undamped

Natura l

Frequency

/-Ma - Mq ü

" o (121)

Es t imates of th e s t a b i l i t y d e r i v a t i v e s for steady l eve l f l l gn t at 1 5 0 knots are

M q

it

m v j

7 6 , 0 0 0 Ib /rad - 1 , 5 0 0 lb / Inch - 11 2 , 0 0 0 f t - lb / rad - 1 5 , 6 0 0 f t - lb-sec/rad - 1 , 3 0 0 f t - lb / lnch 11 , 0 0 0 s lug-f t 2

4 7 , 0 0 0 lb-sec

T he r e su l t ing response parameters are

K -6 . 7 8 deg/sec 2 per Inch

4 « 4.37 sec" 1 T C

Wo0 . 4 3 c 3 .52 rad/sec

T he long i tud ina l control Inputs and p i tch angle responses for Test 414, B l i p s 1 and 2 (Figures 2 0 4 and 71) , have been

u t i l i zed for

the cor re la t ion of

dynamic character is t ics . I n these tests th e control inputs were (for convenience) represented by s ing le impu lse inputs in i t i a ted at zero t ime. They are -1 inch-second for B l i p 1 (Figure i ; 0 4 ) and -2 i nch- seconds for B l i p 2 (F igure 71). he actual pulse durat ions are fa i r ly long compared to the sys tem response period of 2 .0 seconds. or th is reason, an al ternate input represen- ta t ion for Test 414, B l i p 2, has a l so been considered. I n this case tw o success ive steps ar e used; a - 1 . 5 - inc h step at t « 0 i n combina t ion w i t h a 2 . 0 - l n c h step i n i t i a t e d at t 1 second.

3 6 4

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365

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Considering a unit impulse Input ( 1 inch-second), the Laplace transform of the pitch response equation, i n degree units. I s

e ( s ) - -6.78 (3-^.37) ( 1 2 ? ) s[ s 2 + 2( 0 J i30)(3.52)s + (3.52) 2 ]

and the inverse or time solution i s

e = -2 .38 + 2 .58 e - l - 5 2 t sin(3 . l8 t+1.97) ( 1 2 3 )

F or a unit step function, the response equation i s

. , 6.78(s-^. 3 7 ) _ . h ( s

' s2 [ s 2 + 2(0 .n30)(3 .52)s+(3 .52) 2 ] ^

and the corresponding time solution i s

6 = 0.0H 2 . 3 8 t - 0.me' x ' t

si n (3.18 t + 3 . 0 f i )

These solutions have been applied with the Inputs described previously. They are compared t o the recorded responses i n Figures 20 4 and 1 , I t i s seen that for the impulse-input, comparisons are satisfactory for approximately the first 3 seconds, part icular ly considering the corrective stick motion present i n each blip, but the first portions of the pitch rates are overestimated. This i s a n expected result because of the pulse durations mentioned. n ex - amination of the combination-step comparison for Test 4 1 ' ) , B l i p 2 (Figure 71), shows good correlation fo r the first portion of th e pitch angle trace b u t - includes a n unexplain- able delay i n the recorded pitch angle peak of about 1 second. he overal l analysis indicates that i f th n precise stick motion were generated a s an input t o the theoretical system model, th e correlation accuracy would be within 1 degree throughout.

Lateral Dynamics

A theoretical analysis of the pure lateral dynamic response simi lar t o that for the 50-knot case has been made for tne 150-knot tests. I n these runs the stick was pulsed la teral ly i n alternate directions b u t t h e attendant roll angle excursions were not rapidly attenuated, a s i n t h e 50-knot case. This occurred a s a result of excitat ion of the Dutch roll mode wnich i s essential ly nonexistent a t speeds near hover. Consequently, the consideration of tae

3 6 6

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and ts nverse s

* 5.Ml-e* 2' m ) (1^0) Two la tera l tes t r e sponses were s tud ied . They are Test

*1«, Bl ips 3 and t (Figures 98 and 99) , In which the fo rmer i s adequa te ly r epresen ted by a -0 .9 i nch- second impu l se at sero t ime and a correct ive impul se of 1.7 i nch- second t ak ing p lace at 0.75 second , and the l a t t e r input is a 1.1 i nch- second impul se at zero t ime fo l lower by a correct ive impulse of -1 .1 i nch- second app l i ed at 0.5 second. The to ta l r esponse due to both Inpu t s was ca l cu la ted fo r each case us ing the above so lu t ion and super impos ing the i nd iv idua l re sponses weigh ted by thel: respect ive impu l se magn i tudes .

These r esponses are shown in the p lo t s of the measured ro l l re sponses . It i s seen that the cor re la t ion is q u i t e sa t i s f ac to ry f o r t h « * i n i t i a l ro l l responses .

The ver i f i ed con t ro l power and damping levels i nd ica te a s teady-s ta te ro l l ra te of 15 «f t degrees per second per inch o f s t i c k mot ion . In a d d i t i o n the cor respond ing he l ix

ang le o r g^ a lue fo r fu l l con t ro l is 0 .080 , which i s a ® favorable maneuwerab i l l ty level . i t h regard to response t ime, the s t eady s t a t e is reached , essen t i a l ly, in 3 t ime co.istants o r in

3 T 5 7 3 1 1.3 seconds (131)

which is a su ff ic ien t ly rap id response,

3 6 8

r

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(C.

Piasecki Aircraft Corporation

Philadelphia, Pennsylvania

1*. ncrOWT tKCUNiTT CLAIOrtCATlO«

Unclassified

1» 6«OU»

1 ■K^OMT TITLC

16H-1A Flight est Research rogram

1(TTP* »lr

Final Report

»I mnc r'.lumtwm *m*9m)

% AOTKOKH- f A<..i rn m * . mUm» mlifl, ••(

J. H . Goldberg

D. N. Mayers

L. V. Tompkins v mtromr O«TK

August 968

T«. TOTAL

39 8 Tfc. MO. or ncrs

10

•*. COMTMACT OR «MANT M O-

DA 44-177-^MC-154(T) ». rnojlc T HO

Task F162203Al43 l

•A. OHI«INATOI*** RC^ONT NUA«»CM(A»

USAAVLABS Technical Report 67-58

M. OTM«« ««»OTT NOI» fi4jV

10 ntTMOUTIOM TATCMCHT

This document has been approved or public elease and ale;

its distribution s nlimited.

II. •U^n.UaCMTAItT MOT CS 11 - SROMaOnm« MILITANT ACTIVITY

U. S. Army Aviation Materiel Laboratoriei

Fort Eustis, Virginia

1» ASkTMACT

Thi» eport resents he esults f a light est rogram onducted by he

Piasecki Aircraft Corporation on he iasecki Model 6H-1A haft-driven ompound

helicopter. The main omponents f he Model 6H-1A aircraft ncluded: n

H-21 articulated otor, a wing with ontrollable laperons nd hrouded usher

propeller. The aircraft was owered by ingle T58-8 urbine ngine. Aircreit

performance, flying qualities, vibration and ift istibution between main otor

and wing were nvestigated over evel light peed ange of o 67 knots nd dive peeds p o 95 knots.