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RESEARCH MEMORANDUM

INDUCTION SYSTEM CHARACTERISTICS AND ENGINE SURGE

OCCURRENCE FOR TWO FIGHTER-TYPE AIRPLANES

By Terry J. Larson, George M. Thomas,and Donald R. Bellman

High-Speed Flight StationEdwards, Cal.if.

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NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

WASHINGTONMay 26, 1958

,-

-

IlJ

uNCIAsSiFIED

NACA RM H58c14

NATIONAL ADVISORY COMMIT’I’EEFOR AERONAUTICS

RE3EARCH MEMORANDUM

INDUCTION SYSTEM CHARACTERISTICS AND ENGINE SURGE

OCCURRENCE FOR TWO FIGHTER-’IY?E AIRPLANES

By Terry J. Larson, George M. Thomas,and Donald R. Bellman

An investigation was conducted to measure and to compare the total-pressure recovery and distortion characteristics at the compressor faceof two single-place fighter-type airplanes with similar two-spool turbo-jet engines, but with dissimilar inlets. Airplane A has a single normal-shock nose inlet; airplane B has two engines and triangular-shaped inletslocated in the wing roots. In addition, data are presented for enginesurge occurence of these two airplanes and also for a third single-engineairplane having two semicircular-shaped side inlets.

The total-pressure recovery was relatively independent of angle ofattack and mass-flow ratio for both airplanes except for a significantdecrease in pressure recovery with angle of attack for airplane B at thehighest Mach numbers tested. The root-mean-square total-pressure distor-tion decreased slightly with angle of attack and increased slightly withmass-flow ratio for airplane A. For airplane B the distortion increasedwith angle of attack and decreased slightly with mass-flow ratio, partic-ularly at the higher altitudes. Altitude effects on distortion were notedonly for airplane B.

Engine compressor surges were encountered primarily in the regionof high altitude and low inlet total pressure, as indicated by wind-tunnel tests. At lower altitudes and at higher inlet total pressure,surges were encountered under conditions of high total-pressure dis-tortion, particularly circumferential distortion.

INTRODUCTION

Air-flow characteristics of air inlets and ducts of jet airplaneshave a significant effect on the performance of the airplanes. Of these

UNCIASSIFIED

..

A

D

2

characteristics pressure

.

c’”%&%&&&@ NACA RM H5&14

recovery is important, but flow distortion issometimes more important. A large degree of flow distortion can limitengine performance by causing surging in the engine compressor; that is,a stalling of all compressor stages with a rapid fluctuation of air flow.

The air-flow characteristics of a given inlet can vary with alti-tude, Mach number, inlet-flow angles, engine speed, and other parameters.Since it is difficult to exactly simulate flight conditions in a windtunnel, a flight investigation was made at the NACA High-Speed FlightStation at Edwards, Ca.lif.,to measure inlet-flow distortion and total-pressure recovery at the compressor face on three fighter- or interceptor-type airplanes with similar turbojet engines. A part of this investiga-tion was devoted to the determination of surge characteristics of theseairplanes. Since compressor surges can be damaging to the airplane, thenumber of surges encountered was necessarily limited. However, enoughdata were obtained to present a brief comparison of the surge character-istics of the three airplanes. Because the data from the third airplanehave been reported previously in reference 1, only the surge data forthis airplane sre included in the present paper for comparison.

The pressure-recovery data were obtained primarily over a Mach num-ber range from 0.8 to 1.4 and at altitudes between 22,000 feet and42,OOO feet. Compressor surges were encountered at Mach numbers from0.6 to 1.5 and at altitudes between 30,000

SYMBOLS

cross-sectional mea, sq ft

feet and 52,000 feet.

circumferential distortion factor,

( )()( )P’m~ - P’min e- e+— 100, percent

P’av = 180

hp pressure altitude, ft

M Mach number

m/moDuct mass flow

mass-flow ratio,POV@i

N engine inner-rotor speed, rpm

N/~~c normalized engine inner-rotor speed, rpm

NACA RM H5&14 CONFIIENTM

n number of total-pressure probes

P static pressure, lb/sq ft

P’ total pressure, lb/sq ft

P’a~ average total pressure oflb/sq ft

the five total-pressure rakes,

P ‘~~~ highest of the averaged pressures of the five total-pressure rakes, lb/sq ft

P ‘min lowest of the averaged pressurespressure rakes, lb/sq ft

R Reynolds number based on mininn.un

s

T’

v

Wa

‘a rec

x

a

t5c

e

e=

diameter

root-mesm-squsre

/

of the five total-

area equivalent circle

—\

(~1)$distortion, 100, percentn- 1

inlet air total temperature, %

velocity, ft/sec

air-flow rate, lb/see

normalized air flow, lb/see

distance from inlet, in.

angle of attack, deg

P’cpressure normalizing factor,

Sea-level static pressure

compressor-face circumferential

temperature normalizing factor,

station, deg

T’c

3

Sea-level static temperature

CONFIDENTIAL

4

8-

CONFIDENTIAL NACA RM H5&14

circumferential angle subtended by the largest singlyconnected sector of the inlet annulus, having a totalpressure less than the average (see diagram)

e+ circumferential angle subtended by the largest singlyconnected sector of the inlet annulus, having a totalpressure ~eater than the average (see diagram)

P’

u 360°e

air density, slugs/cu ft

compressor-face total-pressure distortion factor,

Subscripts:

o free stream

av average

compressor face

inlet

local

i

1

CONFIDENTIAL

AIRIWINES

NACA R.MH58C14 CONFIDENTIAL 5

The three airplanes are single-place fighter- or interceptor-typeairplanes powered by similar two-spool turbojet engines with after-burners. The following are details of the airplanes and their inletand duct systems.

Airplane A

The engine of airplane A is supplied air through a single normal-shock-type nose inlet. Figures 1 and 2 are photographs of the airplaneand the nose inlet, respectively, and figure 5 is a drawing showing theduct contours and area distribution. The inlet has a lip radius of about0.25 inch, and the diffuser has an equivalent conical expansion angle(included wall angle of a frustrum of a cone having the same length andinlet and exit areas) of 0.37°. Detailed dimensions and other physicalcharacteristics of the airplane are presented in reference 2.

Airplane B

Airplane B is powered by two engines to which air is supplied bytrismgular-shaped inlets located in the wing roots. Each inlet suppliesair to one engine. Figures 4 and 5 are photographs of the airplane andthe inlet, respectively, and figure 6 is a drawing showing the duct con-tours and area distribution. The inlets have elliptical lips with aminimum radius of about 0.3 inch, and the diffusers have an equivalentconical expansion angle of 1.720. Detailed dimensions and other physicalcharacteristics of the airplane are presented in reference 3.

Airplane C

Air is supplied to the engine of airplsme C by two semicircular-shaped side inlets which converge ahead of the compressor face. Fig-ures 7 and 8 are photographs of the airplane and the inlet, respectively,and the duct contours and mea distribution are shown in figure 9. Eachinlet has a lip radius of 0.25 inch, and the diffuser has an equivalentconical expsnsion angle of 1.07°. Additional details on the airplane,duct, and duct instrumentation sre given in reference 1.

CONl?lDENTIAL

6 CONI?IDENTIAL NACA RM H58C14

INSTRUMENTATION ACCUWKY

Each airplane duct was instrumented primarily at the compressorface where five radial rakes for measuring total pressures were installed.Figure 10 shows the circumferential locations of these rakes as well asthe radial positions of the probes for airplanes A and B.

The compressor-face probe pressures were measured by NACA mechanical-optical manometers which determine the difference between the probe andthe reference pressures within *5 pounds per square foot. The flow-distortion values sre based on differential pressures and, hence, wereaccurate to f5 pounds per square foot. The total-pressure recovery atthe compressor face required the measurement of a reference pressure.For airplane A this measurement was made with an absolute manometer cellhaving an accuracy of t20 pounds per squsre foot; for the other two air-planes the reference pressure was determined by a more sensitive method,allowing an accuracy of t5 pounds per square foot.

For all three airplanes Mach number accuracy is believed to be withintO.01 at subsonic and supersonic speeds and within tO.02 at transonicspeeds. Angle of attack was measured to an accuracy of about 0.5°.

RESULTS AND DISCUSSION

The pressure-recovery data were obtained primarily over a Mach num-ber range from 0.8 to 1.4 and at altitudes between 22,000 feet and42,000 feet. Compressor-surge data were obtained at Mach numbers from0.6 to 1.5 and at altitudes between 30,000 feet and 52,0 0 feet. These

2data correspond to a Reynolds number range from 1.4 x 10 to 9.2X 106, .based on minimum-area equivalent circle diameter.

Compressor-Face Total-Pressure Surveys

Contours for every 5 percent of the total-pressure recovery areshown for the compressor face for airplame A in figure 11 for a range ofangle of attack at Mach numbers of about 0.80, 1.0, and 1.4. A compas-sion of the contour plots shows that the distortion is small and, evi-dently, is not effected by either angle of attack or Mach number for theranges presented. The total-pressure recoveries are relatively high,with the lowest recoveries occurring at the bottom of the duct for Machnumbers of 0.8 and 1.0.

Figure 12 presents the same data as shown in figure 11 in a differ-ent., more objective manner. The average total-pressure recovery of all

CONFIDENTIAL

NACA RM H58C14 CONFIDENTIAL 7

probes on each rake is shown as a circle with a solid line connectingthese values. The radial location of each probe and the circumferentiallocation of each rake are both indicated on the abscissa. The dashedline indicates the average total-pressure recovery of all the rakes.From figfie 12 it csm be seen that angle of attack slightly affectscircumferential distortion, especially nesr the 180° region where thetotal-pressure recovery begins dropping when angle of attack increasesbeyond 8°.

Figure 13 shows the vsriation of total-pressure recoveries andtotal-pressure distortions with mass-flow ratio and angle of attack atthe compressor face of airplane A. For the lower Mach number range andfor mass-flow ratios between 0.90 and 1.00 total-pressure recoverydecreases about 2 percent as the angle of attack increases from about0.5° to over 19°. Fairings are not shown for some of the sets of pointsbecause of the large scatter. This scatter is most likely a result ofthe reference cell, noted in “Instrumentation and Accuracy,” on whichthe computations of pressure recovery and mass-flow ratio depend.Despite the scatter, the data indicate that the pressure recovery ishigh throughout the mass-flow range tested and varies little with mass-flow ratio, indicating that for these tests no choking occurred.

The root-mean-square distortion of the total pressure plottedagainst angle of attack (fig. 13) shows the distortion dropping fromabout 3 percent at an angle of attack near 0.5° to slightly less than2 percent at an angle of attack somewhat greater than 19°. The distor-tion also varies slightly with mass-flow ratio, changing from about1.5 percent to about 3.0 percent as the mass-flow ratio increases from0.65 to 0.99.

Figures 14 and 15 present total-pressure-recovery plots forairplane B comparable to those in figures 11 and 12 for airplane A.Figure 14 shows that angle of attack affects both circumferential andradial distortion at the compressor face. It is evident from figure 15that, similar to airplane A, the lowest total-pressure recoveries areexperienced near the bottom of the duct. The average total-pressurerecoveries at a Mach number near 1.0 vsxy from 0.953 to 0.930 in fig-ure 15(b), corresponding to an sngle-of-attack change from 3.3° to 10.7°.Figure 15(c) shows a similar variation for Mach numbers near 1.4.

Presented in figure 16 are the variations of average total-pressurerecovery and total-pressure distortions with angle of attack and mass-flow ratio at the compressor face of airplane B. For the curve showingtotal-pressure recovery plotted against angle of attack the transonicdata are for mass-flow ratios between 0.9 and 1.0, whereas the supersonicdata are for mass-flow ratios between 0.85 and 0.95. In the transonicrange the total-pressure recovery drops from approximately 0.95 to 0.94

CONFIDENTIAL

coNFIDmIAL NACA RM H58C14

when the angle of attack is increased from 0° to about 11O. The super-sonic total-pressure recovery appears to be more sensitive to angle ofattack. At an angle of attack near 0° the recovery is 0.925, while atan angle of attack of 8° it is approximately 0.86. Very little varia-tion in total-pressure recovery is evidenced as the mass-flow ratioincreases from about 0.88 to about 1.00, indicating unchoked flow in theduct .

Figure 16 shows that the root-mean-square total-pressure distortionincreases with altitude as well as with angle of attack. At an angle ofattack of 8° and at a pressure altitude of 50,000 feet the distortion is6.4 percent, at 40,000 feet it is 5.2 percent, and at 30,000 feet (byextrapolating) it is near 4.4 percent. The root-mean-square distortionsof the total pressures for three pressure altitudes plotted against mass-flow ratio indicate little variation in distortion as the mass-flow ratioincreased from 0.88 to 1.00.

Surges Encountered

Several compressor surges were encountered for each airplane.Table I lists flight and compressor-face conditions that occurred imme-diately prior to the surges. In order to summarize the surge regionsand to help analyze the surges, figure 17 presents pressure altitudeplotted against Mach number immediately prior to surge. NACA Lewisaltitude wind-tunnel tests of a similar engine show that compressorsurge can occur at low Reynolds numbers corresponding to total pressuresless than about 500 pounds per squsre foot for normal bleed-door opera-tion and no distortion. These conditions are shown by the shaded regionin figure 17.

Reference 1 shows that for airplane C the compressor-face distor-tions for surge data are not significantly different in magnitude fromthe distortions for nonsurge data and therefore concludes that distor-tion was not responsible for the surges experienced by this airplane.Reference 1 showed, too, in a msnner similar to that used in figure 17,that good agreement is obtained between wind-tunnel data and full-scaleflight data in establishing the surge region of this particular engine.

As cam be seen in figure 17, two of the three compressor-surgepoints for airplane A also occur at total pressures less than 500 poundsper square foot. However, the surge point shown at a Mach number of0.756 was accompanied by a root-mean-square distortion at the compressorface which is considerably greater than normal for this airplane (fig. 15and table I). Although the total pressure at the compressor face exceeds500 pounds per square foot onlyby 40 pounds per square foot, it isbelieved that the surge was caused by the relatively high distortion;

CONFIDENTIAL

NACA RM H5&14 CONFIDENTIAL 9

this distortion was undoubtedly a result of the combination of highangles of attack (21.20) and sideslip (20°) which were experienced imme-diately prior to compressor surge.

Four of the surges for airplane B that are neither in nor near thesurge boundary occurred at Mach rnuibersgreater than 1.35. Reference 4indicates that above this Mach number a surge problem existed until thelower portion of the leading edge of the boundary-layer splitter platewas cambered toward the fuselage, which prevented flow separation on theplate and reduced boundsry-layer buildup. The airplane of this paperdid not have the cambered splitter plate.

If separation occurred on the plate for these surges, it would beexpected that circumferential distortion would occur at the compressorface. By comparing the values of the root-mean-squsre distortions forsurge conditions (table I) with nonsurge conditions (fig. 16), it isseen that no significant difference existed between root-mean-squaredistortion for surge and root-mean-square distortion for nonsurge.However, the root-mean-square distortion is a distortion index whichgives a mes.nvalue of the overall distortion, without distinction between .radial distortion and circumferential distortion. Therefore, it is pos-sible that, if for a particular case the circumferential distortion is“high,” the root-mean-square value may mask this if the radial distortionis “low.” As seen in table I, the circumferential distortion index D(taken from ref. 5) gives higher circumferential distortion values forevery surge that occurred for Mach numbers greater than 1.35 than forthe surges at lower Mach numbers and at total pressures less than about500 pounds per square foot.

For the other point outside the wind-tunnel surge region forairplane B the circumferential.distortion is higher than for the pointsin, or near, the surge region. This high distortion is believed to bethe result of high angle of attack (10.60) of the airplane prior tosurge.

CONCLUDING REMkRKS

Measurements of total-pressure recovery and distortion at the com-pressor face for the test airplanes having similar two-spool turbojetengines, but with dissimilar inlets, indicate that: The total-pressurerecovery was relatively independent of angle of attack and mass-flowratio for both airplanes, except for a significant decrease in pressurerecovery with angle of attack for airplane B at the highest Mach numberstested. The root-mean-square total-pressure distortion decreased slightlywith angle of attack and increased slightly with mass-flow ratio for

CONFIDENTIAL

10 CONFIDENTIAL NACA RM H58C14

airplane A. For airplane B the distortion increased with singleof attackand decreased slightly with mass-flow ratio, particularly at the higheraltitudes. Altitude effects on distortion were noted only for airplane B.

Several compressor surges were encountered for each airplane pri-marily h the region of high altitude and low inlet total pressure, aswas indicated by wind-tunnel tests. At lower altitudes and at higherinlet total pressures, surges were encountered under conditions ofhigh total-pressure distortion, particularly circumferential distortion.

High-Speed Flight Station,National Advisory Committee for Aeronautics,

Edwards, Calif., February 25, 1958.

REFERENCES

1. Saltzman, Edwin J.: Flight-Determined Induction-System and SurgeCharacteristics of the YF-102 Airplane With a Two-Spool TurbojetEngine. NACA RM H57C22, 1957.

2. Matranga, Gene J., and Peele, James R.: Flight-Determined StaticLateral Stability and Control Characteristics of a Swept-WingFighter Airplane to a Mach Number of 1.39. NACA RMH57A16, 1957.

3. Capasso, Vincent N., Jr., and Schalk, Louis W.: F-101 Phase IVStability and Control Test. AFFTC TR 56-14, Air Research andDevelopment Cormnsmd,U.S. Air Force, Jd.y 1956.

4. Pavlick, J. F., and Valor, N. H.: Model F-1OIA - A Summary of theEngine Stall Problem and Its Elimination. Contract No. AF33(600)-8743,Rep. No. 4545, McDonnell Aircraft Corp., Jan. 5, 1956.

5. Alford, J. S.: Inlet Flow Distortion Index. GER-1404, GeneralElectric Co. Presented at the International Days of AeronauticalSciences, Paris, France, May 27 to 29, 1957.

CONFIDENTIAL

NACA RM H5&14 11

confident

12 CONFIDENTIAL NACA RM H5&14

ho

CONFIDENTIAL

NACA RM H5W14 CONFIDENTIAL 13

CONFIDENTIAL

CONFIDENTIAL NACA RM H58C14

In

I

3aJ.->

21-

—

3al.->

4.—U)

CONI?IDENTIAL

NACA RM H5&14 CONFIDENTIAL 15

Pi

f+-d

CONFIDENTIAL

16 NACA RM H58C14

Pi

%o

1

A

Corn?n)mm

I

NACA RM H58C14 17

Duct area

(one side),sq in.

1,000

900

800

700

600

5000 10 20 30 40 50 60 70 80 90 100 110

(a) Longitudinal variation of

x, in.

duct cross-sectional area.

Vane

Side view

(b) Geometric duct characteristics.

Figure 6.- Geometric duct characteristics and longitudinal variation ofduct cross-sectional area for airplane B.

CONI?IDENTIAL

18 comlDmIAL NACA RM H58C14

coIuTDENTIAL

NACA RM H58C14 CONFIDENTIAL 19

Figure 8.- E-2760Photograph showing right inlet of airplane C.

CONFIDENTIAL

20 CONFIDENTIAL NACA RM H5&X4

-,

Duct aresq in.

400’

380

360 ‘

a340

$20

300

280 ~

260

2400 20 40 60 80 100 120 140 160 180 200 220 240 260

Duct station, in.

(a) Longitudinal variation of duct cross-sectional area.

f

Top view

Side view

(b) Geometric duct characteristics.

Figure 9.- Geometric duct characteristics and longitudinal variation ofduct cross-sectional area for airplane C.

NACA RN H5&14 coNFIDmIAL

Accessory

21

fairing

e = 300”

e= 245°

e = 180°

Flush static~

e = 299°

e = 245°

e = 185°

(a) Rake @ramgement ofcompressor facelooking aft.

ane A

95° :

static

(b) Details of rskes.

;a)m

●

w

i

II.

ir guide

Figure 10.- Compressor-face instrumentation for airplanes A and B.

CONFIDENTIAL

22 CONFIDENTIAL NACA RM H58C14

1.

[“1.CQ

_@

a = 2.2”m/iuo . 0.66

a = 10.2”m/y = O.n

a = 4.50m/Is . o.@Q

1.

~

a = 8.0”m/in. . 0.69

a = 12.4”In/m.. 0.72

(a) M = 0.80; hp = 22,000 feet.

Figure 11.. Total-pressure-recove~ contours at the compressor face withangle of attack for airplane A.

CONFIDENTIAL

NACA RM H58C14 coNFIDmIAL 23

a = 2.1”m/m. = 0.93

a - 6.8”ml% o. gk

a . 11.1”m/m. . 0.92

1.CQ

(c2J1.CO

.95

‘“ \ ‘“%.9>

@)

.95

.95

“95>5

Q 1.

.95“9%---.95

a . 14.0”rufmo. 0.90

(b) M = 1.0; hp = 40,000 feet.

Figure 11.- Continued.

CONFIDENTIAL

24 CONFIDENTIAL NACA RM H’38C14

a = O.1°

mlmo= O*B

.95

~

..95

a = 3.9°mi~ . 0.86

(c) M=l.4;

Figure 11.

.95

.95

Q95

.95

a = 2.3”m/me.= 0.87

.95

● 95

~~ .95

.95

a = 5.8”m/m. . O.t%

‘P = 33)500 feet.

- Concluded.

CONFIDENTIAL

NACA RM H5&14 CONFIDENI’IAL 25

0 = 2.20,rnho = 0.66

Figure 12.-

a = 4.50

1.0P’c~ .9

1.0P’c

--

r 0 .9’

h~ = 8.00

dun = 0.69

-L .~ -, --—— ——P’cP’0 .9’

a = 10.2”

A1.0P’cq

.9

.-, .a = J..z.+-

1.0 —P’c i-

-.90 60 I20 180 240 300 360

8, deg

(a) M = 0.80; hp = 22,000 feet.

Circumferential and radial total-pressure-recovery profilesfor different angles of attack for airplane A.

26 CONFIDENTIAL NACA RM H58C14

P’cP’()

‘“”PT--TT

a = 0.6°

In k——— -1

I \l I 11 l’- ! I.9 L I I I I I I I I I I

1.0 ——= —

.9

.9

~ = 6.8.A M/n. = 0.94

_—————Al 0

1.0 .- A h———— —tL.— Q& ––~ — ——

.9’

K1.0 _—— —.

.9

a = 8.bomfmo = 0.94

A@=11.1.

M/n- = 0.92

1.0

.9

a = 14.00raho = 0.93

J’ r--/

—k.

.90 60 I20 180 240 300 360

8, deg

(b) M=l.O; hp=40,000 feet.

Figure 12.- Continued.

CONFIDENTIAL

NACA RM H58C14 CONFIDENTIAL 27

a = 0.10M/m. = 0.88

I .0P‘c /g= — —. + —— fi-– —q I

.9

din; = 6X7l.o -

P’cq ,9

ia= 3.0

IQ& = O. 61.0 .

P’c ~~ —— —— — & “ ——0 .9’

~ = 5.80,

1.0d% = 0.86

P’cT-— — —. * —

q.90

60 I20 180 240 300 360

e, deg

(c) M = 1.40; hp = 33,500 feet.

Figure 12. - Concluded.

28

m

83..00r+

InU3

o

[

.& g

m

coNFIDmIAL

rIefi!ll

CO *

m,

u).

:

(9

N

al

*

o

NACA RM H58C14

v-id

I.

L’

. coNFmmTIAL

NACA RM H58C14 coNl?lDENTIAL 29

1.00A5

~

.95

1.00.0 u1.00

a . 0.7”r+. = O .78

~

6’ 1.00.90

.95

.90

-95

a . 2.7”I@. . 0.82

~:,Gy:.83

= 0.9; hp = 26,500 feet.(a) M

Figure 14.- Total-pressure-recovery contours at the compressor face withangle of attack for airplane B.

CONFIDENTIAL

30 CONFIDENTIAL NACA RM H5&lk

CZ=3.3.m/m. - I.co

1.00

.95

~

.95 .9

1.00

u = 5.1”m/m. . 0.9/3

.90

a= 7.4”m/y . 0 .98

a . 10.7”

m/iQo = o .96

(b) M = 1.05; hp = 42,oOO feet.

Figure 14.- Continued.

CONFIDENTIAL

NACA RM H5&14 CONFIDENIYAL 31

a - 2.0”1(+10 . 0.87

a- 3.8”rufmo= 0.85

85

\

~

.95

.93

\.854

.85

.90

.80

.90

_Q

.95

.90

/%j

~85

m 80-

a . 6.1”r+to . o.e2

.95 .90

~

. 85—.95

.85 0

\ .85.85

k.

a - 7.9°rm/~ . o.&

(C) M=l.40; hp=39,000 feet to

a . 8.5”+0 a o. 82

41,000 feet.

Figure 14.- Concluded.

CONFIDENTIAL

32 NACA RM H5&14

1.0

.9

.8

I .0

,9

.8

I.0

= 0.7”m/m~ = 0.78

“

P’c - ,9pi-

0

.80

m————.T.—60

r——a = 2.7°

a = 4.2°m/m.= 0.83

/ \— ~ “— —

I20 180 240

6, deg

(a) M = 0.9; hP = 26,5oo feet.

E—..,—..——.

II‘J.——

300 3,

Figure 15. - Circumferential and radial total-pressure-recoveryprofiles for different angles of attack for airplane B.

CONFIDENTIAL

IJNACA RM H’58C14

1.0

P’~ ,9P’ o

.8

1.0

.8

a = 3.3°IdaQ ‘ 1.00

a = 5.10m/m. = 0.98

lJ

r ‘ “– ~’,<. ..— .

I

= 7.40dm~ = 0.98

1.0’

a = 8~5°

a = 10.7°

1!0

d% =“0.96

I

A

z

P’c .9— — —

~ —

.8060 120 180 240 300 360

33

(b) M=l.05; hp=42,000 feet.

Figure 15. - Continued.

COIWIDENTIAL

.s

P’cP’(-J .9

%-

CONT’IDENTIAL

Iii!—@

— —

NACA RM H58C14

m—a = 3.8°

RI/mO = 0.85

1.0 ——

P’c .9~

o

.8+

a =6.1°

1.0”IAIO = 0.82

P’~ ,9q

.8

a = 7.9°rnho = 0.82

I .0

P’~ .9

~

.8

I .0

B!LE-

...—

P’~ .9 ?q

.8060 120

.. .—-— -

18/3 240 300 3(

e, deg

(C) M=l.40; hp= 59,000 feet to 41,000 feet.

Figure 15.- Concluded.

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;0

NACA RM H5&14 CONFIDENTIAL

I

35

CONFIDENTIAL

CONFIDENTIAL NACA RM H58C14

hp, f

“1

Figure

32

❑

.4 .6 .8

17. - Pressure altitude andairplanes

I .0M

Mach number atA, B, and C.

AirplaneOA

J-L

a

1.2

incidence

•1

I .4 1,6

of surge for

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