operational features of the langley lunar landing research facility

8/8/2019 Operational Features of the Langley Lunar Landing Research Facility

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NASA T ECHN I C A L NO T E

OPERATIONAL FEATURES OFTHE LANGLEY LUNAR LANDINGRESEARCH FACILITYby Thomas C. O'Bryan a n d Do n a ld E. HewesLangley Research CenterLangley Station, Hampton, Vu.

N A T I O N A L A E R O N A U T I C S A N D S P AC E A D M I N I S T R A T I O N W A S H I N G T O N , D . C . F E B R U A R Y 1 9 6 7


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TECH LIBRARY KAFB, NM

0130b8lNASA T N D-3828

OPERATIONALFEATURESOFTHELANGLEYLUNARLANDINGRESEARCH FACILITY

By Thomas C. O'Bryan and Donald E. HewesLangley R esea rch Center

Langley Station, Hampton, Va.

NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONFor sale by the Clearinghouse for Federal Scientific and Technical Information

Springfield, Virginia 22151 - Price $2.00


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CONTENTS

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 SYMBOLS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 DESCRIPTION OF LUNAR-GRAVITY SIMULATION TECHNIQUE . . . . . . . . . . 4 RESEARCH VEHICLE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

Configuration Features . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 Pilots' Flight- Control Syste ms . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 Fuel Supply System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13 Automatic Balancing System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

GANTRY AND BRIDGE CRANE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16 Gantry Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16 Bridgeunit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17 Dollyunit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19 Operational Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

CONTROL ROOM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22 RESEARCH DATA RECORDING . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24 TEST OPERATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

Preflight Checkout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26 Flight Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27 Postflight and Emergency Operations . . . . . . . . . . . . . . . . . . . . . . . . 27

SYSmM TEST RESULTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 Step-Input Response . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 Flight-Maneuver Response . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30 Observations on Facility O perations . . . . . . . . . . . . . . . . . . . . . . . . 34

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35 REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36

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OPERATIONAL FEATURES OF THE LANGLEY LUNAR LANDINGRESEARCH FACILITY

By Thomas C. O'Bryan and Donald E. HewesLangley Research Center

SUMMARYThe design fea tur es of the Langley lunar landing res ea rc h facility (LLRF) are

described. Thi s facility provides the capability fo r flying a piloted vehicle i n a simulated lunar-gravity field and is presently being used to obtain detailed information relative to the flight dynamics of the Apollo lunar module. Operational procedures andresu lts of a typical flight te st a r e discussed to show the manner in which the facility isutilized. Some of the operational charac teri stic s of the servocontrolled s yste ms whichproduce the simulated lunar gravitational field a r e illustra ted.

Over 150 flight-test operations have been performed to date with 9 rese arch pilotsand astronauts, who have all reported the sensations of actual fre e flight during the te stoperations. Approximately 2 minutes of sustained flight a r e possible by use of the hydrogen peroxide main rock ets in the vehicle. This time capability exceeds the time require dfo r the Apollo lunar module to accomplish its terminal maneuver o r to touch down on thelunar surface from about 150 feet (45.7meters). Continuous monitoring of the se rv o-sy stem s has indicated sa tis fac tory simulation of the lunar gravity , and pilots of thevehicle repor t negligible effects of the visual cues afforded by the facility s tru ctu re.

INTRODUCTIONThe successful perform ance of the touchdown phase of the Apollo lunar module and

of other future manned vehicles intended for lunar and plane tary exploration depends to agreat extent on provis ion of sati sfac tory pilot handling qualiti es. Because present -dayflight experience with manned spacecra ft is limited to orb ital vehicles, the only basis forestablishing handling-quality cr it er ia fo r surface-landing vehicles is the experiencegained with earthbound ai rc ra ft and with seve ral fixed- and moving-base simu lato rs inwhich computers, simulated cabins, and optically generated visual refe renc es are utilized. Recently, two new sim ulat ors, providing actua l flight experience with operationalvehicle s, have become available to assist in establishing and refining the spacecr aftcrit eria . One of these simu lato rs is the Lunar Landing Rese arc h Vehicle (LLRV) basedat Edwards Flight Rese arch Center (s ee ref. 1) and the other is the lunar landing


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re se arc h facility (LLRF) at the Langley Research Center shown in figure 1 . The operational flight envelopes fo r the se two researc h tools are determined by the feat ure s uniqueto each design, but the two simulators are complementary to each other so as to providethe means fo r corre lating actual flight-test experience. Th is is a descriptive rep ort ofthe LL RF covering the lunar-gravi ty simulation technique, the detail design of the var ioucomponents, and the ope ration al aspect s of the Langley facility. A descr iption of apiloted model of the lunar-gravity simulation syst em used t o dete rmine the feasibility ofthe lunar gravitational simulation system f o r the LL RF is presented in reference 2.

L-65-3002Figure 1.- Aerial photograph of the Langley lunar landing research facility as seen -looking toward the east.

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SYMBOLS

Measurem ents fo r the dimensional quantities presented herein we re originallytaken in the U.S. Customary System of Units but a r e presented als o in the InternationalSystem (SI). Details concerning the us e of SI and conversion fact ors relating the twosystems a r e given in reference 3.K1,K2,K3,K4,K5,q,K7 rate and rate-integral feedback gainsS differential operator

H202 hydrogen peroxide

N 2 gaseous nitrogenX,Y,Z vehicle-body-oriented longitudinal, lateral, and vertic al distance,

respectively, feet (meters)

x,y,z down-range, cross-range, and vertical distance, respectively, feet (meters)k,$,k vehicle-body-oriented longitudinal, l ate ral , and ver tic al velocity,

respectively, feet/second (meters/s eca)

X,Y, down-range, cros s-range , and ve rti cal velocity, respectively, feet/second(meters/ second)

.. .. ..X,Y ,z vehicle-body-oriented longitudinal, lateral, and vertical acceleration,

respectively, feet/second2 (meters/seca).. .. ..x,y,z down-range, cr os s- range, and vert ical acceleration, respectively,f eet/second (meters/sec)e angle of pitch, degB angular velocity i n pitch, deg/sec

e angular acceleration in pitch, deg/sec2Subscript:max maximum

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DESCIUPTION OF LUNAR-GRAVITY SIMULATION TECHNIQUE

A basic capability that the lunar landing re se ar ch facility provides is that a pilotis permitted to maneuver the vehicle, shown in figure 2, by means of main lifting andattitude control rockets, anywhere within the li mi ts of the facili ty while under the influence of a simula ted lunar grav itation al field. Simulation of the gravity field is achievedby employing an overhead suspension system, which provides a vertical lifting force equto 5/6 of the vehic le weight by means of cab les acting through the cen ter of gra vity of thvehicle so as to cancel effectively all but 1/6 of the gravit ation fo rce of the ear th. Thecables are attached to the veh icle through a gimbal sy stem which prov ides freedom ofmotion in pitch, roll, and yaw. The upper ends of the cabl es a r e attached to a hoist uniton the traveling bridge cran e shown in figure 3 . This hoist employs a servocontrolledhydraulic d rive syste m which automatically dri ve s the cable s up and down in respon seto the vertica l motions of the vehicle generated by pilot cont rol of the ma in lifting rocket

L-65-4818Figure 2.- Research vehicle k i n g flown d uri ng simulated lunar-gravity operation.

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Load cells in the suspension system provide a signal proportional to the tension in th ecables for controlling the servo-drive unit.

L-66- 1690Figure 3.- View of traveling bridge crane mounted on gantry structure.To maintain vertical alinement of the support cables as the vehicle moves about

horizontally in resp onse t o pilot pitch and rol l commands, a traveling bridge cr an e andan underslung dolly follows the vehicle automatically and stays directly over the vehicleat all times. These units a r e driven by servocontrolled hydraulic syste ms ( simil ar tothat employed in the verti cal hoist unit). The down-range and cro ss- ran ge motions arecontrolled by dolly-mounted cable-angle se ns ors that det ect angu lar deviations of thecable from the vertic al and sepa rat e these deviations into components in the down-rangeand cross- range directions.

The detailed descriptio ns of the various components of t h i s facility a r e given insubsequent sections of this report.

Two operational modes of th e simulation technique were developed; the first mode,which does not actively utilize main lifting rockets, is used primarily for pilot indoctrination and exploratory te s t s of new combinations of con trol- system pa ra me te rs , and thesecond mode which employs operat ion of the ma in ro ck ets is used f or final evaluation ofspecific conditions. Elimination of the main roc kets f or the first mode, hereinafterreferred to as the simulation mode, pe rmi ts all the onboard fuel to be utilized fo r theattitude control rockets. Continuous operation in exce ss of 20 minutes is achieved by

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operation in the first mode, in contras t to the app roxima te 2-minute duration when themain rockets are in use for the second mode of operation . This second mode is hereinafter re fe r red to as the operational mode. The flight tim e of 2 minutes in this mode is iexce ss of that re quired fo r the Apollo lunar module to accomplish the terminal or touchdown phase of its landing maneuve r fr om about 150 feet (45.7 meters).

For the simu lation mode, thr ust of the lifting rock ets is simulated by generatingan elec trica l signal proportional to the commanded thru st and using a potentiometer connected to the throttle control lever. This signal is then added to the signal from th e loadce ll s which normally command a cabl e tension equa l to 5/6 the weight of the vehicle. Inthis manner, the vertical-hoist unit is commanded t o produce an additional for ce on thevehicle through the cab les close ly approximating the ver tic al component of th e rock etthr us t commanded by the pilot. The accele ratio n normally produced by the horizontalcomponent of th e main rockets when the vehicle is at a pitched or rolled attitude is produced in the abs ence of the rocket t hru st by adding, to the cable-angle-sensor signals,electrical biasing signals proportioned to the pitch and ro ll angles so as to command acable angle offset fro m the vert ica l equal to about 1/6 t he corresponding vehicle attitudeangle. In or de r to produc e the commanded offset cable angle, the bridge or dolly mustaccelerate at a leve l consistent with the angular attitude of the vehicle. Because of thependulous action of the cab les, the approximate na tu re of the additional command sig na lsto the s erv o drive systems, and the requirement fo r the vertical drive system to acceler at e the entire m as s of the vehicle, some er r o r s in the response of the vehicle a r e associated with this mode. However, the advantage of a greatly increased flight durationexceeds the disadvantage of small dynamic e r ro r s for the indoctrinary and exploratorypha ses of the facility operatio n.

The operational mode provides g re at er re ali sm of flight by eliminating the dynamice r r o r s associated with the simulation mode of operation and als o by adding the audio cu esof the main rockets' firi ng and the constraint of limited fuel and flight time. The lat tertwo factors produce a pilot stress level which is gr ea te r than that imposed by the simulation mode and probably is much clo ser t o that imposed by an actual lunar landing.

RESEARCH VEHICLE

The piloted vehicle, shown close up in fig ure 4, was designed as a rese arch tool sothat many of the sy stem p ar am et er s and design feature s could be varied as pa rt of ageneral re se ar ch prog ram on the handling qualities of lunar flight and landing vehicles.Control-system adjustm ents with minimum modification ar e provided by the us e of analogcompu ters mounted on the vehicle. A list of the physical ch ara cte ris tic s of th e vehicleand the ranges of variable system parameters is given in table I.

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Figure 4.- Closeup of research vehicle. L-66-2144

TABLE I.- PHYSICAL CHARACTERISTICS AND DESIGN FEATURES O F THE L LR F VEHICLETotal weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 000 lbm (5443 kg)Fuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90 percent H202Maximum fuel capacity . . . . . . . . . . . . . . . . . . . . . . . . . . . 3100 lb (1406 kg)Maximum nitrogen capacity . . . . . . . . . . 19 f t 3 at 3000 ps i (0.54 m3 at 20 684 kN/m2)Moment of inertia:

Pi tch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3540 slug-ft2 (4800 kg-ma)Roll (including gimbal support) . . . . . . . . . . . . . . . . . 4770 slug-ft2 (6467 kg-ma)Yaw (including gimbal sup port) . . . . . . . . . . . . . . . . . 4750 slug-ft2 (6440 kg-mz)

Maximum height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.75 f t (4.5 m)Pilo t's eye level above ground . . . . . . . . . . . . . . . . . . . . . . . . 13.17 f t (4.0 m)Landing gear :

S p a n . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.5ft (3.5m)Shock strut type . . . . . . . . . . . . . . . . . . . . . . . . Air-oil (honeycomb, backup)Shock stru t str oke. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 f t (0.3 m)Maximum load factor, g-units (earth) . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

Main lifting thru st . . . . . . . . . . . . . . . . . . . . 600 to 6000 lbf (2.669 to 26.689 kN)Control moments:Pitch . . . . . . . . . . . . . . . . . . . . . . . . . 100 t o 2000 ft -l b (136 to 2712 m-N)

Roll . . . . . . . . . . . . . . . . . . . . . . . . . . 128 to 2560 ft-l b (174 to 3471 m-N)Yaw . . . . . . . . . . . . . . . . . . . . . . . . . . 135 to 1350 ft-lb (183 to 1830 m-N)

Maximum attitude angles:Pitch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . *30Roll . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . *30Yaw . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . *360

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The res ear ch vehicle is somewhat s ma ll er than the Apollo lunar module and thepilot is sitting down ra th er than standing up; however, the linear and angular acce ler ations produced by the main and the attitude rocke ts are comparable. Consequently, thevehicle per mi ts an accur ate duplication of lunar module flight char acte rist ics.

Configuration FeaturesThe basic s tru ctu re for the vehicle is a box-like frame 4 feet (1.2 m) deep by

8 feet (2.4 m) squar e to which are externally attached the four landing gear legs, thepilots' compartment, the gimbal suspension assemblies, and the propulsion system.The gimbal assemb lies house the pitch gimbal bearings, which c ar ry the cable loadsinto the vehicle str uct ure and allow the vehicle to rotate fr eely in pitch within the lim itsof 30' pitch up and 30 pitch down. Roll and yaw freedom is provided by the whiffletree,which is discussed subsequently.

Each landing-gear leg employs a main shock strut, such as that used on helicoptersattached by a pivot joint to the co rne r of the vehicle fra me . The stru t is stabilized bytwo articula ted a r m s attached to the bottom of the vehicle fram e. A 1-foot-diameter(0.3 m) aluminum landing pad is attached at the apex of the landing-gear leg by means ofa bal l-socket joint, which allows the pad to aline itself with the slope of the te rr a in whenground contact is made. The pad is spring loaded, as shown in figure 5, so as to rotateon the ball-socket joint as the vehicle leav es or makes contact with the ground, whereby

Figure 5.- Closeup view of landing-gear pad and slipper. L- 66-2192

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elec tric al switches, one on each leg, are caused to open and close. The purpose of theswitching action is to short circuit the integrating amplifiers in the various s er vo syste ms (discussed in subsequent sections) so as to insur e stable operation of the s er vo sys t em s with the vehicle on the ground. Also shown in figure 5 is a 1.5-foot-diameter(0.46 m) stee l pad that was incorporated as a "slipper" so as to reduce high frictionalfo rc es between the aluminum pad and the concrete landing surface. These high fo rc esprevent the shock strut fro m functioning properly. The shock-strut assembly incorpora te s aluminum honeycomb elemen ts that begin t o cr ush when the vehicle vert ical velocity exceeds 10 feet/second (3.05 m/sec). Sufficient stroke and energy absorption isprovided by the honeycomb to permit landing with a verti ca l velocity of about17 feet/second (5.18 m/sec) without exceeding the vehicle design loading.

The pilots' compartment consis ts of the front portion of a helicopter modified to bemounted directly on the vehicle fra me . A plot of the field of view from the right-handseat is shown in figure 6. The instru ment panel shown in figu re 7 is considered to be abas ic s e t of inst rume nts fo r lunar landing and is arranged for primary pilot control fromthe right-hand sea t. These inst rume nts display roll, pitch, and yaw attitudes and angularrat es, and the coordinate line ar positions and velociti es rela tive to the landing site . Athree-gimbal stable platform and body-axis-mounted rate gyros provide driving signal sfor t hese displays.

Other quantities pertaining to vehicle systems operation displayed on the panelinclude fuel flow and fuel quantity, main-engine manifold pressure, and electrical supplyvoltages. An oxygen breathin g system employing standard mili tary a ir cr af t oxygensupply equipment is provided f o r the occupants of the pilots' compartment. Thi s syst emis provided as a safety measure to prevent the pilot from inhaling the noxious productsof the rocke t motor exhaust ari sin g from the decomposition of hydrogen peroxide fuel.A circulating cold water suit is provided f or the pilots to wear on days when the weatherconditions result i n cockpit temp erat ures , inside the closed plexiglass bubble, above about850 F. Electric al power is supplied in the vehicle during flight by a group of nickel-cadmium ba tte rie s located ahead of the pilots' compartment.

Pilots' Flight- Control SystemsThe pilots' p rima ry flight controls a r e shown in figure 7. The main lifting engine

throttle lever, formerly used as the helicopter collective pitch control, is operated bythe left hand, either f ro m the right-hand or left-hand pilot's seat. The re a r e currentl ytwo readily interchangeable attitude contro l arra ngements provided in the vehicle. Thefirst is the two-axis side-arm controller, shown at the right-hand seat , which is used inconjunction with the foot-operated yaw pedals. The second arra ngement is the three-axis controller, shown at the left-hand seat, and is sim ila r to the controller used in the

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m

6c i lert ical

40UP

'EI20 Position of pilot's eyes

.-d in r ight-hand seat5L0- 0.-3c-.-Uc2 20Down

40

60

8o120 100 80 60 40 20 0 20 40 60 80 100 120

Figure 6.- Plot of fiel d of view fro m rig ht- han d seat of pilots' compartment.

e Figure 7.- Pil ot's in st ru me nt panel. L-65-4522.110


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lun ar module. The yaw pedals are locked when the three-axis controller is being usedand pilot's yaw-con trol input consi st s of a twisting w r i s t motion about the vertical axisof the control stick. Clos er deta ils of these two control arrangem ents are shown infigures 8 and 9.

The pitch-control sys tem of the vehicle is illustrated diagrammatically in figu re 10. This system is typical for all th re e axes. Attitude-control torq ues are generated by twenty 25-pound to 125-pound (111 to 556 N) thrust rocket motors locatedaround the periphery of the vehicle fram e. The multiplicity of motors, eight fo r pitch,eight for roll, and four for yaw, was used to provide flexibility in the re se ar ch operations. These motors can be ground adjusted over the thrust range to provide a widerang e of angular accel erat ions about each of the control axes. The mo tor s ar e operatedin a n on-off o r bang-bang fashion by solenoid-operated valves, which obtain their firin gsign als when the s um of th e pilot-input control s igna ls and the selec ted feedback signalsfed into the electronic relay switch unit exceeds a given voltage level. This voltagelevel, which deter min es the sys tem dead band, is pr es et by the pilot on a control consolein the pilo ts' compartment. An additional dead band of + loof controller stick trav el isprovided for each axis of control. Diode li m it er s prevent the signal from the cont roll erpassing until the voltage output equivalent to lo of displacement is exceeded. A deadband is required in a system of t h i s type to prevent excessive fuel consumption as aresul t of limit-cycl e operation and inadvertent pilot control inputs.

L-65-4520.1Figure 8.- View of two-axis controller (pitch and roll control) and rudder (yaw) pedals.

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L-65-4519.1Figure 9.- View of three-axis controller for combined pitch, roll, and yaw control.

Pilot inputY-Y p$ Vehicle motion

Attitudecontroller d RateI gyro

Pitch-attitude control system

Ballvalvelh ro t t lecontroller

Throttle con trol systemFigure 10.- Schematic diagrams of pitch-attitude and the throttle control systems forresearch vehicle.

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The feedback signals co nsist of the a ngular rate and the inte gra l of the angularrate and are used to provide different types of attitude-control sy st em s to be evaluatedin the re se ar ch program. The gains of the s ignals can be individually adjusted by thepilot. With both signal ga ins set at zero, the system operates as an open-loopacceleration-command system which provides a constant and continuous angular acceleration as long as the pilot's control stick is deflected outside the dead band. With therate feedback gain adjusted to a finite value, the system operates as a rate-commandsyste m in which displacement of the control stick resu lts in a corresponding angularvelocity after the attitude mot ors have fire d momentarily. This syste m limits the maximum angular rate that the vehicle can achieve and also damps ang ular distur bance s whichproduce velocities outside the equivalent rate dead band of the sys tem .

Addition of the rate -int egr al signa l produ ces an attitude-command sys tem whichgenerates a change in vehicle attitude directly proportional t o displa cemen t of the pilo t'shand controller. The attitude engines f ir e until the attitude e r r o r resulting fro m thestic k displacement and the ensuing angular r at es have been reduced to less than the corresponding ra te and attitude dead bands.

Control of the main-rocket th rust through a range of about 600 to 6000 poundsforce (2.669 to 26.689 kN) is achieved by means of the throttl e contro l syste m, al sodiagramed in figure 10. A pneumatic boost actuator positions a ball valve in the fuelsupply line immediately ahead of t he main rocket s in re sponse to th e pilot's throttlemovements. A push-pull cable fr om the throttle oper ates the actuator through a nonlin ear cam used to compensate the nonlinear flow char acte rist ics of the ball valve therebyachieving a linear throttle-position- to-rocket- thrust relationship.

Fuel Supply SystemA schematic of the fuel sy ste m of th e vehicle is given in figu re 11. Pressur ized

nitrogen gas at 3000 psi (20 648 m/m2) is stored in six titanium pressure tanks equippedwith a pressure relief valve set at 3250 psi (22 400 kN/m2). Th is gas is used to pressu riz e the four interconnected stainless-steel fuel tanks at 675 ps i (4654 kN/m2) througha dome-loaded regulator-vent valve and a hand-operated isolation valve. A pressurerelief valve set at 700 psi (4826 kN/m2) is used to prevent overpr essur izatio n of thetanks in ca se of the nitrogen regulator malfunction.

The fuel tanks have a to tal capacity of about 3100 pounds (14062 kg) of 90 pe rcen thydrogen peroxide which is used to operate the main lifting rockets, the attitude-controlsyste m, and the pitch-trim system . Total fuel flow is measured in the main fuel line byan in-line fr ee- rot or flowmeter employing magnetic pickup. The fuel-flow signa l isintegrated electronically to determ ine the quantity of fuel used. Pneumatically operatedfail-safe valves are used i n the fuel lines supplying the main engines and attitude-control

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-- --

sys tem to perm it fuel cutoff eith er by the pilot or from the control room in cases Ofemergency and to in su re safety while the crew is handling the vehicle. Hand-operatedball valves located just upstre am of the main mo tor s are utilized as an additional safetyfactor for vehicle handling pr io r and subsequent to a flight.

ARelief

EmergencyAttitu de rockets

Mai n rocket I i systemsTrim8 rockets 8

Figure 11.- Schematic diagram of the fuel system for the research vehicle.

P r e s s u r e of the fuel used in the pitch, roll, and yaw bran ches of the attitude-controlsystem is regulated independently by dome-loaded regulators. Fir ing of the attitudeengines is controlled by in-line solenoid-operated va lve s immedia tely ahead of each attitude rocket. A number of hand-operated valv es a r e located in the low spots of the system to perm it complete drainage of th e fuel fro m the vehicle.

The fuel system utilizes stain less stee l throughout to provide compatibility withthe hydrogen peroxide fuel and the piping and valves a r e dimensioned to allow flow r a te sof 50 pounds per second (22.7 kg/sec) to the main engines.

System pre ssu res are displayed on the pilot's instrument panel and also at themechanic's panel attached to the fr am e at the r e a r of the vehicle. Adjustment of all thepress ure regulators is made at the mechanic' s panel and operation of the emergencyvalves is controlled by the pilot.

Automatic Balancing SystemBecause the m a s s of fuel used during a given flight represents about 25 perc ent of

the total ma s s of the vehicle, t he problem of center-of-grav ity shif t caused by fuel consumption and fuel tr an sf er within the tan ks while the vehicle is pitched and rolled is14


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---extreme ly criti cal. The seve rity of thi s problem is illustrated by the fac t that themomen ts generated by the shifting fuel can exceed the test-con trol moments in som ecases by a fa ctor of 10. The magnitude of this problem is grea ter i n an earth simulationof lunar gravi ty and is minimized i n the actual luna r landing situation inasmuch as thelun ar weight of t he fuel will be very much le ss and th er e will be essentially no tendencyof the fuel to shift with th e attitude changes. Several s tep s were taken in the originaldesign of the vehicle and in subsequent vehicle modifications to re duce the effects ofcenter- of-gravity shift.

For the purpose of minimizing the vert ical center-of-gravity shift due to fuel consumption, the four cylindrical fuel tanks were located as close to the vehicle cente r ofgravity as possible; in addition, an automatic ballast mechanism, which moves a weightvertically with a compensative motion, was developed. This ballast mechanism, locatedon the fr am e of t he vehicle, con sis ts of a 250-pound (113.4 kg) ballast box riding onvertical trac ks and driven by a servocontrolled system, which is positioned in responseto a pulsed el ectrical signal derived from the fuel flow measuring system.

The tanks, alined in the fore-and-aft direction , we re compartmented and heavilybaffled to minimize fuel tran sf er from one end to the other as the vehicle is pitched.The se precautions to limi t the horizontal s hift of fuel we re not adequate for pitchingmotions because of the rela tive ly long duration of lar ge pitch ang les used in flight maneuvering; the gravity grad ient within the tanks allowed only the fuel in the higher end of thetanks to flow to the rock et motors. Consequently, the rocket-powered pitch-trim s yste m,shown schema tically in figu re 12, was installed on the vehicle. (See als o fig. 4.) Thissys tem, by producing an opposing moment with the appro pria te pa ir, of boom-mountedrockets, compensates for the moment due to fuel shift. The thrus t levels of the rocketpa i r s are modulated from an idling thr us t of about 30 pounds (133 newtons) to the maxi mum of 125 pounds (556 newtons) by the action of ser vocon tro lle d thro ttling ball valves.The servos for the two separate valves obtain their positioning signals from an angularaccelerometer alined with the pitch axis of the vehicle . Stabili ty and minimum system

Pitch-up Ball Pitch-upservo valve rocketsA f ir ing switch d z t ' iner t iaPitch-downYTh alve Pitch-downI-

Figure 12.- Schematic diagram of th e automatic pitch -trim system.

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. . . . . .. . . . . ...... .

e r r o r ar e provided by combining the pitch-acceleration signal with the integra l of t hepitch signal. A uniform and nearly linear variation from z er o to maximum trimmin gmoment in either di rection is achieved by moving only one throt tle valve at a time andholding the other at its idl e position. Thi s action avoids having to shut off one pa ir ofrockets while the other is turned on, so that an undesi rable discontinuity in rocket o peration is avoided.

In ord er to prevent cance llation of pilot-applied p itch con tro ls by this system, acalibrated biasing signal corresponding to the commanded acceleration is added wheneverthe pitch-attitude control r ocke ts are fired . Since the action of this system is always toseek either a zero or a commanded pitch-acceleration condition, the system h as theadditional desirable feature of automatically compensating fo r extran eous momentsacting on the vehicle such as those produced by winds and pilot body motions.

GANTRY AND BRIDGE CRANEThat pa rt of the facility which s upports the vehic le and follows its linear motions

consists of a gantry structure and a traveling crane. The traveling cra ne consists of abridge str uct ure that tra vel s the length of the gantry and an underslung dolly that tr av el sthe width of the bridge.

Gantry StructureThe gantry structure, shown in figure 1, is oriented in the east-west direction and

is composed of truss elem ents arr ange d with four s et s of inclined legs that provide adequate cle ara nce fo r any pendulous motion of the vehicle which could develop. An elevato r enclosed in a shaft at th e east end provides a cc es s to the overhead equipment, andcatwalks permit all pa rt s of the s tru ctu re to be inspected at close range.

The structure is about 240 feet (73 m) high with two 400-foot (121.9-m) long tr ac ks72 feet (22 m) apa rt at the 220-foot (67-m) elevation. Each tr ac k is a flat surface18 inc hes (46 cm) wide coated with an epoxy paint embedded with carborundum particlesto provide good tract ion of the bridge rubber -tire d d riv e wheels. The str uct ure ispainted in an alternating pattern or orange and white co lor s at about 40-foot (122-m)inter vals to meet U.S. government regulations relative to aerial st ructur es . Some of theorange sections on lower portions of the leg s were omitted to minimize the possibilitythat the known dis tan ces of the alte rnat ing color bands might provide the vehicle pilot swith optical cues to their altitude. The operational a re a underneath the gantry shown infigure 2 is paved with a 30-foot (9 m) wide conc rete s tr ip to minimize jet-blast effectsand fuel- spillage problems.

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Bridge UnitThe bridge unit shown in figure 3 is composed of a tubular box-frame structure,

10 feet by 10 feet by 62 feet (3 by 3 by 19 m), supported at both ends by four-wheeledtru ck as sem bl ies utilizing standard 52-inch-diameter (1.3 m) heavy-duty tru ck ti re s.Total span of the bridge is 72 feet (21.9 m) between center lin es of the two wheel as semblies, which are driven independently by the servocontrolled hydraulic drive systemsmounted in each end of the bridge stru ctur e. The two lower span memb ers of the bridgestructure form a tr ac k which suppor ts the dolly unit and permit s it to move in the cr oss -range direction.

Each bridge driv e unit is composed of a constant-speed induction electric motor,rated as 250 horsepower (186 500 watts) driving a variabl e stro ke hydraulic pump whichis in turn directly coupled to two constant-displacement hydraulic mo tors of the samesize and rating as the pump. (See fig. 13.) Each motor is geared t o two of the drivewheels through a 4.5 to 1 pinion-bull g ea r power train. The hydraulic-fluid circuit is aclosed-loop sys tem with di re ct connections between inlet and outlet port s of t he pumpand motors. Thi s syst em util izes 3-inch (7.6-cm) sta inle ss-s teel pipe that is pressuretest ed to 9000 psi (62 053 kN/m2). Maximum operating p re ssu re is approximately3000 psi (20 682 kN/m2) regulated by pr es su re relief valves in the hydraulic mot ors.This hydraulic circu it is supercharged by mean s of an auxiliary pump driven by theelectric motor. An elevated 30-gallon (0.09 m3) sump is used t o provide makeup fluidand to rece ive servovalve, pump, and motor-cas e leakage.

L T o ervoamplifier and driveunit on other side of bridgeFigure 13.- Schematic diagram of bridge drive unit showing basic electronicand hydraulic circuits.

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Control of each hydraulic d rive unit is provided by a servo-operated hydraulicvalve, which positions the drive-pump st ro ke r and thereby regu late s the pump flow output and, consequently, the bridge 's speed and direction. An additional set of servo-operated safety valves, located in each hydraulic motor, bypass the flow fro m each motorthrough internally mounted pre ss ur e- relief va lves to provide a closely regulated brakingforce for either a normal or an emergency braking operation. During thi s action thepump stroker is also automatically returned t o its zero-stroke position and the pump-outlet port is ported directly to the inlet port to unload th e pump.

Electric power fo r the induction mot ors is ca rri ed from the gantry to the bridge bymeans of bus b a r s mounted parallel to and directly over the bridge t ra ck on one s ide ofthe gantry and a trolley -collec tor assembly mounted on the sam e side of the bridge.Elec tric s igna ls for control of the servo-operated hydraulic valves as well as severalother functions ar e c arr ied to the bridge from the gantry by a long trai ling cable whichis dragged along by the bridge in a tray beneath the track as shown in the left-hand sideof the photograph i n figure 3 .

The two independent drive sys tem s at opposite ends of the bridge a r e adjusted tosh ar e the driv e loads equally by proper balancing of the elec tronic ampl ifie rs associa tedwith the circui try of each sy stem shown in figure 13 . The command signal fo r both sys tems during lunar-gravity operations is derived fro m cable-angle se ns or s located on thedolly. The cable-angle signal is summed along with the first and second int egral of th issignal pr io r to being attenuated by a variable potentiometer attached to the cable drum onthe dolly. The function of thi s potentiometer is to var y the system gain as a function ofthe cable length to insu re system stability at all cable lengths. The potentiometer outputsignal is passed on through a preamplifier to the power amplifie rs which driv e the bridgehydraulic servocontrol valves.

When rem ote opera tion of the bridge is required for positioning the bridge andvehicle before and after each te st operation, an open-loop control signa l is genera ted byhand-operated potentiometer controls at the ground-control station.

Several mechanical, hydraulic, and electrical feat ures a r e incorporated to providesafety in operation of the bridge unit. A cr o ss shaft linking the two side s of the bridgetogether is used to tran sm it unsym metrica l loads in the event of fa ilu re of one of thedriv e sys tems . Lateral motion of the bridge is limited by a se t of two spring-loadedguide rol ler s at each end of the bridge. These ro ll e rs engage the si de s of the curbingand allow trave l of about 3/4 inch (1.9 cm). This small lateral trav el is required tominimize stru ctu ral loads on the bridge and to allow for the variations in the tra ck gagecaused by manufacturing and erection tole rances of the gantry stru cture . The bridge isrestrained from jumping the track by the action of the cross-shaft assembly which gripsthe top plate of the curbing and by the guide rol le rs which rid e under th is s am e plate.18


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The built-in hydraulic relief valves i n each drive motor prevent excessive pres su r es and corresponding stru ctur al loads from building up in the system. Several pr es su re switches throughout the s ystem detect lo ss of hydraulic pr es su re and automaticallyshut down the syst em in a safe manner. The hydraulic moto rs function as normal brakingdevices to halt the motion of the bridge, but additional braking is provided by means ofexpander-tube frictio n br ak es mounted within each of the drive wheels. Braking is normally applied in a cascading sequence which begins with the pump stro ke r being drivento ze ro str oke by grounding the servo-power amplifie r. At the same time, action of thefriction brake s is initiated but a needle valve in the brake lines cau ses the full brakingfor ce to build up ove r about 2 seconds. At about 1.5 seconds the safety valves arereleased so that the pump is completely isolated from the moto rs and each motor canfunction as an independent hydraulic brake. This particu lar fo rm of braking sequenceinsures a uniform application of the braking fo rc es fro m the different braking devicesand minimizes the possibility of excessive braking for ces due to simultaneous operationof the devices o r lo ss of braking action due to fail ure of any one device.

All electric al rela ys and hydraulic valves a r e se t up so as to operate fail-safe i nthe event of lo ss of el ect ric power. Also limi t switches are installed at the extremetrav el limit s of the pump s tr ok er s to initiate emergency braking in c as e of a hard-overtype of elec tric al or hydraulic failure. Similarly, limit switches a r e provided near theextr emes of the bridge travel. An additional braking action is provided by mechanicalbuffers, located at each end of the gantry t ra ck to absorb re sidu al bridge velocities up to7 feet/second (2.1 m/sec).

Dolly U n i tThe dolly is a box-frame stru ctur e about 10 feet squ are (3 m) and 4 feet high

(1.2 m), suspended on the t ra ck beneath the bridge unit by m eans of four se ts of ste elro ll er s mounted at the co rn er s of the dolly. The vert ical hoist syst em andSee fig. 3 .)the cros s-ra nge drive system a r e mounted on the dolly f ra me and utilize a common250-horsepower (186 500-watt) el ec tr ic motor driving the two separate servocontrolledvari able displacement pumps of each system. The prin cipl es of operation for these twosyst ems, including the action of the brak ing syste ms, are sim ila r to those for the bridgeunit.

The power tr ai n for the cro ss-r ange driv e system co nsis ts of the dri ve pump onthe north s ide of th e dolly and a matching 35-horsepower (26 110-watt) hydraulic mot orgeared t o the center of a driv e shaft interconnecting the two si de s of the dolly. A pinionat each end of the shaft engages racks which are an inte gral pa rt of the two track s onwhich the dolly rides. A fric tion brak e acting directly on the interconnecting dri ve shaftprovides the safety backup braking fo r the cross-r ange driv e syst em.

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The power tra in for the vertical-h oist driv e co nsis ts of the hydraulic pump andmoto rs identical with those used in the bridge drive system. In this case, however, thepump is connected to thr ee moto rs in parallel and these m otors in turn a r e geared 4t o 1 to a common driv e ge ar in tegral with the cable drum. A friction brake a cts directlyon the end of the cab le.d rum opposite from the dri ve gear. The cable dru m is groovedand provided with a level-wind mechan ism and guid es to preve nt fouling of the two250-foot-long (76.2 m), 7/8-inch-diameter (2.2 cm ) cables. A third cable, carryingabout 60 separ ate elec trica l and instrumentation wire s, is wound on a separate cabledrum driven by chain from the main cable drum. Each of the tw o main cables passe sthrough a guide in the floor of t he dolly and through the previously mentioned 5-foot-long(1.5-m) cabl e-angle- sens or as se mb lies mounted to the bottom of the dolly, as shown infigure 14. The function of th e tube is to sen se the angular deviations of th e cable s fro mthe vertical in the two orthogonal planes alined down range and c r o s s range. An electrical signal proportiona l t o the deviation in each plane is generate d by r ota ry potentiometers geared t o the gimballed mounts supporting the sens or tubes.

Figure 14.- Closeup of two cable-angle sensors mounted to bottom of dolly. L-66-2145.1

As shown in figure 15, the lower ends of the cabl es a r e terminated by swagedfittings attached to the vehicle gimbal support assem bly, which cons ists of a spreaderba r housing a yaw swivel unit, a pivoted rocker beam, re fe rr ed to as a whiffletree, andtwo ve rti ca l support st ru ts with self-alining bearings . The function of this asse mbly isto transmit the cable for ce to the vehicle so that the forc e ac ts through the vehicle c ente r20


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Figure 15.- Details of vehicle gimbal support assembly. L-65-5200.1

of gravity but permits the vehicle to rotate freely about all axes within specific limits.The limits selected for operation a r e *30 in pitch and ro ll and *360 in yaw.

The load cells used to generate the ele ctric al signals fo r the servocontrolledvert ical hoist system a r e mounted in struts. The outputs of the two cells are summedelectrically so as to obtain the total cable force and to elim inate any effe cts of rollingmoments. The resultant signal is conditioned so that the vertica l driv e system will maintain a tension in the support cables equal to 5/6 the weight of the vehicle. (See fig. 16 . )

Operational LimitsThe flight envelope in this facility is illustrated in figure 17 . The vehicle can be

flown anywhere within a corr idor 360 feet (110 m) in the down-range X-direction, 42 fee t(12.8 m) cros swis e in the Y-direction, and 180 fee t (54.9 m) vertically in the Z-direction.The serv odrive sy st em s can follow the vehicle at velocitie s up to about 25 feet/second(7.6 m/sec) in the X-direc tion, 8 feet/second (2.4 m/sec) in the Y-direc tion, and1 5 feet/second (4.6 m/sec) in the Z-direction.

Compensation is provided f or the drag fo rc e on the vehicle and cables due to avera gewind velocity. There is , however, no compensation for the effect of aerodynamic mom entson the control of vehicle attitude. These moments are of c onc ern only fo r low angularaccelerations (5 deg/sec2 or less) with the vehicle attitude-contro 1syst em; consequently,these tests are generally re st ri ct ed to days when the winds are relatively calm.

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-----

\ ixed displacement hydraul ic motors The bridge crane was designed toand safety valve assemblies (3 ) provide the capability of flight testingveh icle s weighing up to 20 000 pounds(9072 kilograms); however, currentoperational limitations restrict vehicle

Servovalve an dpump strokerf weight to about 1 2 000 pounds(5443 kilograms).CONTROLROOM

Manual Suspension control I cables Operat ion of the fac ilit y is con-(control room)+OP trolled and directed from an observationIns t rumentcable- + room on the second floor of the officeIII ._ . . --- and shop building located near the south-I-

electronics - vehicleFigure 16.- Schematic diagram of vertical drive unit showingbasic electronic and hydra ulic circuits.

Figure 17.- Operational envelope for lunar landingresearch facility.

2 2

wes t corne r of the gantry str uctu re.From this vantage point movements ofthe vehicle, bridge , and dolly a r e viewedby the test director and facility operato rs through la rge observation windows.(See fig. 18.) The room is equipped withcontro ls for manual and automatic operation of the bridge-crane dr ive system s.A number of instrument disp lays indicat e the statu s and perform ance of thesedrive systems as well as of the vehicleitself, and three plot boards a r e used toindicate motions of the system relativeto the safe operating limi ts. Althoughthe system will disengage automaticallyin a safe manner when these limit s a r eexceeded, manual controls for disengagement and subsequent braking action a r eprovided in the control room, as well asin the vehicle, as a safety backupfeature.

Two-way communications are provided throughout the facility so that thetest director is in voice contact at allti me s with the pilot and the operational


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(a ) View looking north. L-647236.1

(b)View looking east. L-64-7237.1Figure 18.- Control-room console.

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I ., , . .....-....-.... .....

crews which s erv ice the facility and provide photographic coverage during the t es toperations.

RESEARCH DATA RECORDING

Several fo rm s of data recording are utilized so as to accumulate and retain thediffe rent type s of information r equi red in the operation of the facility and the evaluationof the flight tes ts. The information is used to determine that the automatic servo-operation of th e facil ity is providing rea lis tic lunar-gravity simulation, to show what thetest pilot does to control the vehicle while performing the test maneuver, and to recordthe vehicle response.

In addition to the previously mentioned control- room plot boards which indicate andreco rd the relation of positions and velocitie s of the bridge, dolly, and hoist, a direct-writing oscillograph is used in the control roo m to obtain time h ist ori es of the cableforce and the cross-range and down-range angles.

Two alternate types of flight re co rd er s a r e available for use on the vehicle duringthe tes ts to obtain time hi stor ies of about thir ty channels of information pertaining topilot inputs, vehicle sy st em s operation, and vehicle response as listed in table 11. Theprimary system is a multiplexed seven-channel magnetic tape-rec order syste m whichper mit s replay of the dat a within a few minutes after the test run is completed so as to beof ass istanc e during the post-flight evaluation of the flight test and debri,efing of the pilot.

TABLE 11.- SUMMARY OF RECORDER DATA CHANNELSONBOARD THE LLRF VEHICLE

Parameter RangeCable force . . . . . . . . . . . . . . . . . . . . 0 to 15 000 lbf (0 to 66 723 N)Fuel pressure. . . . . . . . . . . . . . . . . . 0 to 550 psi (0 to 3792 kN/m2)Throttle stick position . . . . . . . . . . . . . . . . . . . . . . . . . Oo to 41'Firing signal (each axis) . . . . . . . . . . . . . . . . . . . . . . . . On-offStick position (each axis) . . . . . . . . . . . . . . . . . . . . . . . . +looAngular acceleration (pitch and roll only) . . . . . . . . . . . . . k30 deg/sec2Angular velocity . . . . . . . . . . . . . . . . . . . . . . . . . . +50 deg/secAngular position:

Pitch and roll . . . . . . . . . . . . . . . . . . . . . . . . . . . . &30Yaw . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . +360

Bridge velocity . . . . . . . . . . . . . . . . . . . . . +25 ft/ sec (7.62 m/sec)Dolly velocity . . . . . . . . . . . . . . . . . . . . . . k10 ft /sec (3.05 m/sec)Cable angle, range and cross range . . . . . . . . . . . . . . . . . . . . . . +3'

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Each of th e seven channels on the magne tic tape contains five channels of analog inform ation which have been converted t o frequency-modulation signal s and multiplexed p ri or t otaping. Replay of the tapes and recording of the da ta on visua l reco rd s immediately a fte reach flight a r e achieved by us e of a demodulation syst em and the direct-writing oscillograph in the control room.

The second vehicle recording system, which is installed only when the pri mar y sys tem is inoperative, consists of a standard NASA oscillographic flight rec orde r. Theindividual information channels are recorded directly on a 6-inch film strip and are readout by means of semiautomatic data process ing equipment, a process requiring severaldays.

Photographic views of the flight te st s are obtained in the fo rm of magnetic-taperecorded television pictures and 16-milli meter motion-picture film. The televisionrecords are used f or ini tial postflight evaluation and pilot debriefing but a r e notretained. The motion-picture f il ms a r e used for subsequent evaluation and permanentrecords. Both television and film pic tures are obtained from a manually operatedca me ra station just outside the control room at the southwestern c or ne r of the facility.Two additional manned statio ns for motion pictu res a r e located at the southeast co rner, one at the 50-foot (15.2 m) eleva tion and the other at the 150-foot (45.7 m) e levation. An additional rem ote c am er a fo r motion pic ture s of the vehicle is located in thedolly, with the c am er a looking straight down so as to obtain an overhead view of thevehicle relativ e to the ground. Dire ct mea sureme nts of th e vehicle motions can beobtained by m eans of this motion picture and a calibration s ca le painted on the con creteapron over which the vehicle flies.

Additional photographic coverage is obtained inside the vehicle by two fixed,remotely operated motion-picture ca me ra s to obtain reco rds of the pilot's activit ies andfield of vision. One 16-mil lim eter c am er a mounted on the sid e of the ins trum ent paneland looking direc tly a t the pilot' s face re co rd s hi s head and eye motions. The secondcamera is located beside the pilot's head between the two s ea ts and views the instrum entpanel and a portion of the window area. A m i r r o r is located on the panel so that themotions of the pilo t' s r ight hand on th e attitude contr oller can be seen by th i s camera .

Audio reco rdings of the flight t e s t s a r e obtained by use of a tape r eco rde r in thecontrol room connected to the communications system.. This record includes the conver sat ions of the flight di recto r and test pilot as well as the other facility o per ato rs whotake an active part during the time the pilot is in the vehicle. An audio tape record ing isals o made throughout the pilot "debriefing'' ses sion subsequent to each flight operation.

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TEST OPERATIONS

The complete daily test operation is performe d in essentially thr ee phases, compris ing the preflight, flight, and postflight activities . The preflight activ itie s a r e concerned with ground-crew efforts to insur e c or re ct and safe operation of all subsystemsand integrate d units. The flight operation is concerned pr imar ily with the activities ofthe tes t pilot, flight director, and monitoring crew to perfo rm the specific flight-testmaneuv ers in accordance with the t es t plan and within the safe operational limits. Thefinal postflight activities are concerned with sec urit y mea su re s to permi t safe exit ofthe pilot and ground handling of the vehicle fo r the next te st flight o r for sto rag e at theend of the operational day. Some asp ect s of th ese act ivit ies a r e disc usse d in the following sections.

Preflight CheckoutEach morning of a te st day the bridge and dolly units a r e operated remotely from

the control room and tested to all limit positions at typical operational speeds t o demons tr at e the adequacy of all fail -safe equipment. Following the init ial checkout of thevehicle in the hangar, the vehicle is moved to the a pron at th e east end of the gantry 'bymeans of a tr ac to r and special tr ai le r. Connections a r e made to the gimbal suspensionstruts, the interconnecting instrumentation cable, and an umbilical cable which suppliespower to the vehicle until the actual flight operation begins.

After the nitrogen tanks and the hydrogen peroxide tank s a r e loaded, the vehicle israised about a foot (0.3 m) above the apron and final ballastin g is provided to ins ure thatthe vehicle is balanced properly. A prefligh t check of the servo-ope ration of the ve rt icalhoist is then made by rai sin g the vehicle to an altitude of about 150 fee t (45.7 m) wherethe system is placed into automatic operation. This action allows the vehicle to fall asthe s ervo syst em maintains the desi red tension in the cabl es equal to 5/6 of the vehicl e'sweight. The vehicle accel era tes downward with the resulting 1/6 ear th gravity until the15-foot-per-second (4.6 m/sec) velocity lim it of the hydraul ic syste m is reached. Atthis point the system is braked automatically by the limit switches on the pump stroker.

A checkout te st of the bridge and dolly serv osy ste ms is perform ed with the vehiclerest ing on the ground and with the cab les under tension. This test cons ist s of displacingthe bridge and dolly 2 fee t (0.6 m) fro m directly over the vehicle so that a cable angle ineach direction of about 0.5O is produced. The bridge and dolly drive syst ems a r e thenplaced into automatic operation to allow the bridge cra ne to rea line itself over the vehiclewithin about 0.05'.

The pilot e nt er s the vehicle by way of a removable acc es s ladder and establishescommunication with the test director through his standard flight helmet, which also hooks


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into the oxygen breathing syst em, as soon as the servosystem checkout tes ts a r e completed. The pilot and ground crew begin activation of the vehicle flight sy ste ms in accordance with check lists and guidance of the te s t dir ector . The pilot switches over to internal flight batt erie s and the external power umbilical is removed. The propulsion- systempr es su re regu lators a r e then adjusted by the mechanics, and the fuel tanks ar e pressuri zed by the pilot. A walk-around inspection is made by a mechanic fo r fuel or gasleaks, after which the stable platform is activated, and all rockets a r e warmed up bymomentary control inputs by the pilot. The vehicle is then rais ed cl ea r of the ground bymanual operation of the bridge cr an e as the pilot operates the attitude controls to maintain a level attitude. The pilot then activate s the flight-r ecor der syst em and sequentiallyapplies step inputs to each control for a syst em calibration. The vehicle is then movedto the des ired position for the start of the flight tes ts , which may begin with the vehicleon the ground or at some altitude up to about 150 feet (45.7 m), depending on the particularflight-test program.

Flight Te st sThi s flight-te st phase of operation begins with a countdown performed by the flight

director, which results, sequentially, in the safety brakes' being released by the control-room o perator, the flight r ec or de r' s being turned on and the throttle adjusted to approximate hover-thrust position by the pilot, and, finally, the bridge c ra ne 's being put intoautomatic operation by the control-room operator. A t thi s point the pilot is in completecommand of th e vehicle and init iate s the flight maneuvers to ca rr y out a preplanned flightprofile. The flight director and control-room monitoring crew observe the vehicle anddisplay instrumen ts to evaluate the flight maneuvers relati ve to the safe operationallimit s of the facility. In the event that the limi ts a r e approached, the dire ctor eitherwarn s the pilot o r stop s the test, depending on the s ever ity of the situation. Thepilot may a lso stop the tes t i f he becomes aware of some operational problem. N o r mally, however, the flight is carr ied to its conclusion by completing the planned flightmaneuver and landing. The system is stopped or disengaged immediately by the flightdirector after the vehicle comes to a rest on the ground.

Postflight and Emergency OperationsIn the event that the flight is terminated by disengagement of the dri ve system and

subsequent braking, the pendulous swing of the vehicle is permitted to subside prio r tolowering the vehicle to the ground by the manual operation from the control room. Ifthere is a malfunction in the vertical-h oist drive systems, the vehicle c an be lowered bymanipulation of the braking sy stem by eit her the control-room operato r or the cre w on-board the bridge cra ne. A mobile cherry-picker is on standby status fo r use in r a r eemergency situations when the vehicle cannot be lowered to the ground.

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With the vehicle on the ground, the pilot depressurizes the fuel and nitrogen storagetanks, and the ground crew moves i n to flush down the vehicle and ground area with wate rto remove any traces of fuel. A fire-fighting crew and equipment are on standby statusat the facility in the event of fi r e resulting from spilled fuel. The pilot egr es se s andre tu rn s to the operations building f o r a postflight debriefing s ess ion while the vehicleeither is refueled fo r the next flight in about 45 minutes o r is disconnected fro m t hecables and returned to the hangar for stora ge at the completion of the day's operationalactivities.

SYSTEM TEST RESULTS

Some res ul ts of typical operational te st s of the facility a r e given i n figures 19 to24 to illust rate the response of the bridge cra ne systems, first, to step inputs used toest ab lis h optimum gains and, second, to typical pilot-controlled landing maneuve rs of thevehicle used in the pilot-handling re se ar ch program.

Step-Input ResponseThe bridge and dolly re spon ses are si mi la r and ar e illustrated by the typical bridge

test data shown in figure 19 which show the var iat ion s of th e bridge velocity X anddown-range cable angle as a function of t im e following a st ep cable-angle signal equivalent

fVsecm/sec12 13 r

/Whiff letree oscil lations

Time, sec

Figure 19.- Time histories of cable angle and bridge down-range velocity i n response to a 0.4' cable-anglestep input.28


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to about 0.4O. The vehicle is suspended a few feet off the ground with the dolly and hoi stsys tem s inoperative. This type of test dem ons tra tes the ability of th e bridge and dollyto maintain a desired cable angle, near Oo, over the full operational ran ge of t he sys tems.At the initiation of the test, the input signal causes the bridge to a cceler ate forward atabout 0.2 foot/second2 (0.06 m/sec2) , towing the vehicle along so as to produce the commanded cable angle. The oscillographic t r ac e of cable angle in figure 19 shows that th isaverage angle was attained 1.8 secon ds aft er initiation of the st ep input with a subsequent50-percent overshoot of about 0.2O. The period of the pendulous cable oscillation wasabout 8 seconds or approximately half of the natural pendulous-cable period of 15.7 secconds, which occ urs when the bridge servos ystem is inoperative. Following the initia langle overshoot the oscillation is attenuated and the cable angle is maintained toabout io.loo.

Low-amplitude h igher-frequency oscilla tions of 1 cycle per second and higher a r esuperimposed on the pendulous-cable frequency because of the bending o r whipping of thecables. The limiting fac tor fo r gain of the cable-angle signal, which provid es attenuation of the pendulous motion, is the divergence of the bending freq uencie s which a r e higherthan the drive -system natural frequency. There fore, a near-optimum cable-angle gainis achieved when the re is som e evidence of these higher frequencies. Elimination ofcable-angle e r r o r due to vehicle velocity is achieved by us e of the cable-angle inte gra lsignal, but this signal has a destabilizin g effect on the pendulous-cable motion; consequently, the optimum gain of the int eg ral s ignal is determined by the system characteristi cs. The gain of the double-integral signal which eliminates the cable-angle e r r o r dueto vehicle acceleration is also limited for the same basic reason.

The test of the ve rtic al hoist syste m is illustrated in figure 20, which shows theoscillographic tr ac es of v ert ical velocity and cable for ce as the vehicle is allowed to f a l lfr e e under the simulated lunar gravity provided by the servocontrolled hoist. The vehiclefalls until the maximum velocity is reached, at which point the syst em is automaticallydisengaged. The cable-force t ra ce shows the decr eas e in cable tension to 5/6 the vehicleweight as the system is switched into automatic operation. The transient oscillations,primar ily caused by cable stretch, a r e damped; and the force change is essentiallyunaffected as the velocity i nc re as es linearly with time to a maximum of 14.5 fe et pe?second (4.42 m/sec) in 3 seconds. The avera ge acceleration determined fro m the velocity trac e is 5.3 feet/second2 (1.6 m/sec2), which corres pon ds to the lunar-grav ity value.The re su lts of a series of t es ts to determine hoist system cha racteri stics have indicatedthat it has a natura l frequency of approximately 1.5 cyc les pe r second with a dampingratio of 0.5.

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UL

W

60-

40 -

--

al0- 8000 --l0n - Test start -w Velocity l imi t20 4ooo-

o * I I I I I I .. I L * . 1 I0 -

Figure 20.- Time his tori es of vertical velocity measured at ho ist dr um and cable force measured at vehicle in responseto a free-fall test dur in g simulated lunar-gravity operation.

Flight-Maneuver ResponseThe trajectory of a typical landing fro m an altitude of about 100 feet (30 m) and a

horizontal d istance of 180 feet (5 5 m) from the landing spot is shown in figure 21 , whichis an X-Z coordinate plot of the flight path with 5-second t im e int erv als marked alongthe path. Tim e hi sto rie s of the down-range velocity, down-range cable angle, cable force,and vertical velocity fo r this sam e flight test a r e presented in figure 22. The flight wasinitiated from a hovering condition as the pilot elected to genera te a vertical descentprior to pitching over to move forward to the designated landing target. The vehicle w a spitched down to accelerate forward at about 0.3 foot per second pe r second (0.09 m/sec2)from 5 seconds to 27 seconds during which time the cable angle deviated from the vertica lby no mor e than 1/30. For the int erval fro m 27 seco nds to touchdown, the vehicle waspitched up t o provide deceleration from a maximum down-range velocity of 5.8 feet persecond (1.8 m/sec). The cable angle rev er sed direct ion but did not exceed the value30


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ftm 12035

0 sec5

" 0 20 40 60 80 100 120 140 1601 1 , I 1 I I I I --'lm0 10 20 30 40 50 60Down-range distance, X

Figure 21.- Typical flight-pat h profi le of research vehicle perfor ming a land ing maneuver from altitude of about 100 feet(30 meters) and 180 feet (55 meters) away fro m the selected down-range landi ng spot.

of 0.30. The pilot made many cor recti on s of the main lifting engine thru st throughout thelanding maneuver as is indicated by the abrupt changes in the slopes of th e ver tica lvelocity-time-history curve. These changes correspond to ver tica l accelera tions ofabout 3 feet per second pe r second (0.9 me ter/ seca) downward to 1.5 fe et pe r secondper second (0.45 meter/s eca ) upward. These acceleration changes, in turn, represen tve rt ical thr us t changes of about 1200 pounds (5338 N) downward and 600 pounds (2669 N)upward. The time-history curve of cable force rev eals that, as ide from the small cable-stre tch oscillations induced by the ra ther abrupt thrust changes, the cable fo rc e is unaffected by the lar ge vert ical motion of the vehicle. The cable for ce slowly dec re as es inresponse to the vehicle-weight change resulting f rom fue l consumption. The averag erat io of cable fo rce to vehicle weight rema ins constant at 5/6 throughout the flight.Transient oscil lations equivalent to *0.3 lunar rr g? s"occur dur ing flights, as indicated i nfig ure 22; however, the pilots have generally been unable to perceive thes e oscillations.The relation of t hese typical flight motions relat ive t o the operational l imi ts of th efacility are illu stra ted i n fig ures 23 and 24 which are the safety monitoring re co rd s displayed in the control room by two of the plot boa rds d iscu ssed previously. Figure 23shows the altitude-vertical-velocity plot with the superimposed ver tica l operationalboundaries, which denote a maximum descent velocity of 14 feet per second (4.3 m/sec),and a maximum asce nt velocity of 13 feet per second (4.0 m/sec). The slanted portions

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W-

W

--

kN60 1 2 0 0 0 C ~,nit ial weight Final weight\

'

w 40L 8000 c -Ds 2o - 4000

m-0-m ~. . . . . . .: . 4 L LE ols I East-.41 , ' 1 I , -,-L , - j 1 - - 1 I I 1 1- I _ _ - I 1 I 1 !0 5 10 15 20 25 30 35 40 45 50 55 60

Time, secFigu re 22.- Time histories of vertical velocity, cable force, down-range velocity, an d cable angle for la nding maneuver shown in figure 21.

3 2


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c

m 120353O-lo0

-25 - 80U, 203.- 60 -a

15-40

10 -205

1 I I -IJ fVsecOL O20 15 10 5 0 -5 -10 -15' 4I I I . I 0I I -2 --m/sec -4Vertical velocity UP

Figure 23 - Altitude-vertical-velocity plot obtained from control-room plot board showingmotions of typical landing maneuver (figs. 21 an d 22) relative to operational limits offacility.

of the boundaries r ep rese nt the braking and driving capabi lities of the hoist. The slantedsolid line in the "down" sid e of the plot corresponds to the velocity-exceed boundary whichautomatically disengages the system when the limit line is exceeded. The dashed line isused as a warning boundary for the contro l-room plot-board monitor. The flight-testtrace reflects the altitude and velocity changes shown in the previous figure and illustrates that the pilot was able to perform the flight maneuver well within the operationallim its of the syst em.

The plot of down-range velocity as a function of distance down range in figure 24illustrate s the operational limits for the bridge system. These limits do not represen tthe maximum capabilities discussed previously but have been employed to provide extramargin of safety during init ial operation of the facility. The flight-t est t ra ce shows thatmaneuvers si milar to this part icul ar flight can be perf ormed well within these operational limits.

There were no significant cross-ra nge excursions for this test as the pilot wasmaintaining flight prim aril y in the XZ-plane. In other flight test s, particularly thosewhere lateral translation and yawing maneuvers have been performed, the maximumavailable cro ss-r ang e lim its of *21 feet (6.4 m) have been encountered usually after the

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' ft300L I I I I - I I .. I0 20 40 M1 80 l i 0East West

Down-range distanc e

Figure 24.- Variation of down-range velocity with distance obtained from control-room plot board showing motions of typical landing maneuver (figs. 21 an d 22)relative to operational limits of facility.

pilot has initiated attem pts to co rr ec t the overtravel condition. To avoid disengagementof the system and discontinuance of an otherwise sat isfacto ry flight te st when such asituation occurs , a modification to the normal automatic operation has been incorporatedso as to permi t continuation of the flight so long as all sy ste ms a r e operating normally.Thi s modification, which in effect provides a cushioning action, consist s of a se t ofswitches which slow the dolly to a halt without disengaging any of the sys te ms as the dollyapproaches within 2 feet (0.6 m) of the normal ex tre me trav el-l imit switches. Thisaction per mits the pilot to complete the correcti ve actions and bring the vehicle back tothe des ire d flight path and continue with the remaining flight maneuver. In the eventthat the pilot has failed to take a corr ecti ve action and approaches these switches at anexcessi ve speed, the dolly will slow as these switches a r e encountered but will also over-tra ve l and str ik e the extr eme- limi t switches, disengaging the syst em and applying brakingaction.

Observations on Facility OperationsOn the ba si s of nearly a hundred manned flight t es ts in which eight re se ar ch pilots

and astronau ts have been used as pilot test subjects, the following general observationsand commen ts relat ive t o the facility operation can be made:

(1)The servocontrolled bridge, dolly, and vertical- hoist sys tems operate sa tisfactorily t o produce the desi red v erti cal lifting force on the flight vehicle and thereby simulate the effects of lunar gravity on the vehicle flight behavior.

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(2) The pilots, without exception, have repor ted sensa tions of actual fr e e flight anda r e unable to detect artifac ts attributable to the servocontrolled system during normalflight-test operations.

(3) The physical operat ional limi ts of the facili ty have imposed only minor const ra in ts on the pilot's ability to perform landing tasks. The primary constraint ha s beenthe total available cross- range travel of 42 feet (12.8 meters ) which has generallyrequired somewhat m or e attention on the pa rt of the pilot to lateral drif t than might berequired otherwise.

(4) The visual obstruction or cues imposed by the gantry s truc ture appear to havea negligible effect on the pilot 's ability to perfo rm a typical landing task.

(5) Utilization of the simula ted main -thrus t mode of operation is very useful f or theindoctrination of new pilot test subjects and for performing exploratory res ea rc h flightswith untried combinations of control-system pa ra me te rs and vehicle design featur es.

(6) The facility ap pea rs to be practic al fo r continued r es ea rc h into the flightbehavior and pilot handling ch ar ac te ri st ic s of not only the Apollo lunar module but al soother vehicles proposed fo r operation in lunar and other subgravity environments. Withonly minor modifications to the suspension fittings and adjus tment s of the s ervosy stem,the facility can be utilized fo r studies of lu nar roving vehicles, rocke t propulsion backpacks for lunar and space operations, and M a r s o r Venus rocket-propelled landingvehicles.

CONCLUDING REMARKSDesign fea tur es and operational cha rac teri stic s of the lunar landing res ea rc h

facility at the Langley Resear ch Cente r have been discussed. Resul ts of a typical flighttest are included to il lus tra te the capabilities of the facility in simulating lunar gravity.Observations based on rep or ts of nine pilots, including th re e astronauts, from nea rly100 flight te st s indicate this facility to be a practical rese arc h tool for assessing theflight behavior and handling ch ara cter isti cs of space vehicles proposed for operation insubgravity environments.Langley Research Center,

National Aeron autics and Space Administration,Langley Station, Hampton, Va., October 19, 1966,

125- 19-01- 13-23.

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REFERENCES

1. Bellman, Donald R.; and Matranga, Gene J.: Design and Operational Characteristicsof a Lunar Landing Re se arch Vehicle. NASA TN D-3023, 1965.

2. Boisseau, Peter C.; Schade, Robert 0.; Champine, Robert A.; and Elkins, Henry C.: Preliminary Investigation of the Handling Qualities of a Vehicle in a Simulated Lunar Gravita tio nal Field. NASA TN D-2636, 1965.

3. Mechtly, E. A.: The International System of Units - Physical Constants and Conversion Factors. NASA SP-7012, 1964.

36 NASA-Langley, 1967 L-4920


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operational features of the langley lunar landing research facility

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