a compilation of recent research related to the apollo mission

8/6/2019 A Compilation of Recent Research Related to the Apollo Mission

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TECHNICAL MEMORANDUM X-890

A COMPILATION OF RECENT RESEARCHRELATED TO THE APOLLO MISSION

By Langley Research Center StaffLangley Research CenterLangley Station, Hampton, Va.

NATIONAL AERONA UTICS AND SPACE ADM INISTRATION


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CONTENTS

LUNAR ORBITAL ENTRY AND DESCENT

. LUNAR-ORBIT LANDING SITES AND STAY TIMES . . . . . . . . . . . . . . . .By Robert H. Tolson

. RETURN-TO- EARTH CONSIDERATIONS AND STAY-TIME CONSTRAINTS FOR THELUNAR MISSION.1By John P. Gapcynski. A SIMPLE GUIDANCE TECHNIQUE FOR LUNAR LETDOWN. . . . . . . . . . . .1By Richard Reid LUNAR LANDING

. VISUAL SIMULATION STUDY OF LUNAR HOVERING, TRANSLATION, ANDTOUCHDOWN.7By Maxwell W. Goode. LANDING CHARACTERISTICS OF A MODEL OF THE LUNAR EXCURSION MODULE . . . .7By Ulysse J. Blanchard. LUNAR TOUCHDOWN OPERATIONS AS AFFECTED BY ABORT CONSIDERATIONS . 47By Gary P. Beasley. DEVELOPMENT OF A SIMULATOR FOR STUDIES OF SELF-LOCOMOTION OF MANUNDER REDUCED GRAVITY.3By Donald E. Hewes and Amos A. Spady, Jr.TAKE-OFF, RENDEZVOUS, AND DOCKING. SOME ASPECTS OF MAN'S VISUAL CAPABILITIES IN SPACE . . . . . . . . . . .9By Jack E. Pennington. FIXED-BASE GEMINI-AGENA DOCKING SIMULATION ....7By Byron M. Jaquet and Donald R. Riley

. THREE-DEGREE-OF-FREEDOM FIXED-BASE SIMULATION OF PILOT-CONTROLLEDLUNAR TRAJECTORIES FROM LIFT-OFF TO RENDEZVOUS . . . . . . . . . . .9By Charles P. Llewellyn. MANUAL CONTROL OF A LUNAR LAUNCH ..5By Lindsay J. Lina2. AN OPTICAL DEVICE FOR OBTAINING ATTITUDE AND ALTITUDE. . . . . . . .3By Alfred J. Meintel, Jr. i


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13. ANGULAR SEPARATION OF APOLLO AND LUNAR EXCURSION MODULE AT LANDERTOUCHDOWN.01By James L. Williams and L. Keith Barker

1 4 . SIMPLE ABORT SCHEME FOR SYNCHRONOUS ORBITS. . .. . . . ...07By G. Kimball Miller, Jr., and L. Keith Barker1 5 . ABORT CONSIDERATIONS FOR THE LUNAR EXCURSION MODULE . . . . . . . . . .1 5By David B. Middleton1 6 . PENETROMETER TECHNIQUES FOR LUNAR SURFACE EVALUATION. . . . . . . .23By John Locke McCarty, Alfred G. Beswick, and George W. Brooks1 7 . PROPOSED SURVEYOR LANDING EXPERIMENT.31By Sidney A. Batterson

1 8 . RADIO-FREQUENCY SIGNAL ATTENUATION BY PLASMAS OF ROCKET EXHAUSTGASES.............................. 135By Duncan E. McIver, Jr.1 9 . DYNAMIC PENETRATION AND EROSION OF DUST-LIKE MATERIALS IN AVACUUM ENVIRONMENT.45By Leonard V. Clark and Norman S. Land2 0 . VISIBILITY AND DUST EROSION DURING THE LUNAR LANDING. . . . . . . .55By Leonard Roberts

LUNAR RESEARCH FACILITIES

21. LOLA, THE LUNAR ORBIT AND LANDING-APPROACH SIMULATOR . . . . . . . . . 171By William T. Suit and Ralph W. Stone, Jr.22. DESCRIPTION OF A LUNAR -LANDING RESEARCH FAC ILITY . . . . . . . . . . .79By Thomas C. O'Bryan23. RENDEZVOUS DOCKI NG SIMULATOR . . . . . . . . . . . . . . . . . . . . .87By Howard G. Hatch, Jr.24. NEW DYNAMIC RESEARCH FACILITIES SUITABLE FOR SUPPORT OF APOLLO--LUNAR-EXCURSION -MODULE MISSION . . . . . . . . . . . . . . . . . .93By D. William Conner25. LUNAR-GRAVITY SIMULATOR FOR FULL-SCALE IMP ACT TESTING . . . . . . . . . 199By Robert W. Herr 11


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LUNARORBITAL ENTRY

AND DESCENT


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5

1. L UNA R -ORBIT LANDING SITES AND STAY TIMESBy Robert H. Tolson

SUMMARY

The determination of possible lunar-landing sites and the correspondingfor missions in which the lunar-orbit-rendezvous

trajectories, are reviewedindicates that a

INTRODUCTION

In the initial Apollo missions the location of the landing site on the lunarprimary parameters determining the over-

scientific or other reasons; therefore,lunar sites which are accessible for the Apollo

of course, an integral part of the lunar-orbit-rendezvous technique and, ascan be established playl u n a r - l a n d i n g s i t e s . I t i s a s s u m e d

launch from Cape Canaveral and that after aninjection into the transfer trajectory occurs at

480 km with an injection angle of O o . T y p i c a l v a l u e s o f t h e s e t w obecause the resultst o e s t a b l i s h t h e l u n a r o r b i t . T h e

the inclination of thel o n g i t u d e o f t h e a s c e n d i n gline S2.

Although the lunar-landing procedure for the lunar excursion module has notthe lunar excursion module will

t h e c o m m a n d m o d u l e . T h e w i d t h o f t h i s b a n d w i l lmake orbital


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possible lunar-landing sites is directly related to the geometrical characteris-tics of lunar orbits which can be established efficiently and which at the sametime are consistent with overall mission requirements.

GEOMETRICAL PROPERTIES OF LUNAR ORBITS

In order to investigate the possible landing sites it is advantageous firstt o d e t e r m i n e t h e a l l o w a b l e r a n g e o f i a n d 2 c o n s i s t e n t w i t h t y p i c a l c o n -straints on the transfer trajectory. For this purpose, it would be convenientto have explicit relationships between the lunar orbital characteristic and theinjection conditions. Unfortunately, no exact expressions are known and someapproximations are required in order to obtain general analytical information.For this study the earth and moon are assumed to be spherical masses and to movein circular orbits at their mean distance; in addition, a "patched" conic tech-nique (ref. 1) is utilized. This technique is based on the assumptio n that thereis an imaginary sphere centered at the moon with a radius of about 58,000 km.When the vehicle is inside the "sphere of influence," the earth's gravitationaleffects on the vehicle are neglected; and when the vehicle is outside of thesphere, the moon's gravitational forces are neglected. With these assumptions itis possible to obtain explicit relationships between the lunar orbital character-istics of interest here and the transfer trajectory injection conditions. Detailsof this analysis have been reported in reference 2, but a few of the results arementioned herein because they have a direct bearing on the lunar-landing-siteproblem. In figure 2 the sphere of influence is depicted with the moon at itscenter. The latitude q measured from the earth-moon plane and the longitudemeasured from the earth-moon line are used to specify the location of any pointon the sphere of influence. The dashed lines represent the projection onto thesphere of influence of two typical lunar orbits and as before the orientation ofthese planes will be specified by the inclination and nodal position.

A nominal transfer trajectory is defined as one with a specified injectione n e r g y a n d a s p e c i f i e d i n c l i n a t i o n t o t h e e a r t h - m o o n p l a n e I . A n u m b e r o f f r e eparameters at the injection point are still to be considered and if these freeparameters are properly varied, it is in general possible to obtain a wide varietyo f l u n a r o r b i t s , n a m e l y , l u n a r o r b i t s w i t h a w i d e r a n g e o f v a l u e s o f i a n d Q .However, the analysis of reference 2 showed that the orbital planes of all lunarorbits established from a given nominal trajectory pass through a common point onthe sphere of influence, that is, all the orbital planes have a common line ofintersection. This result provides a relationship between the inclination, nodalposition, and the latitude and longitude of the common point of intersection oft h e o r b i t a l p l a n e s, a s fo l l o w s:

tan i =tansin(Q- )Thus, if the energy and inclination of the transfer trajectory are specified, thevalues of ^ and q are fixed; consequently, the inclination and nodal positionare not independent parameters and once one is chosen the other is determined by

(1 )

2


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. v

This common point of intersection is often called the entry point becauseof influence through which the vehicle must pass

the moon's surface; that is, the extended trajectory

ower part of the figure gives th e locus oftrajectory inclinations to the earth-moon

third of this region which

Briefly, what has been shown about the establishment of lunar orbits is thatlunar orbits established

thelatitude and longitude of

a l l o w a b l e v a r i a t i o n i n ^d n as they appear in the relation cannot be changed appreciably by varyingnominal trajectory.

LUNAR-LANDING SITES AND EXPLORATION TIMES

With the foregoing relationships established between the lunar orbital char-by lunar-orbit-rendezvous requirements may be considered.

e s t a b l i s h a l u n a r o r b i t w i t h a n i n c l i n a t i o n i g r e a t e r t h a nthat thel a n d i n g s i t e i s 8 d e g r e e s o u t o f t h e

be a relative motion betweenabout 1 3.2 0 p e r d a y . ( N oteare con-

o f f s e t i s i n c r e a s i n g s o t h a t t h e " r e t u r n t o

3


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P = GO.2 days2)wherecos D8=os i sin ^ - sin 83 )2os 'T sin iAt any time during the exploration period the lunar excursion module can returnt o t h e c o m m a n d m o d u l e w i t h n o t m o r e t h a n a n o f f s e t o f S d e g r e e s r e q u i r e d .Now, the significance of the previously derived relationship between orbitali n c l i n a t i o n a n d n o d a l p o s i t i o n i s r e a l i z e d . F o r i f 8 i s s p e c i f i e d b y t h e p r o -pulsion capabilities of the excursion module, then the latitude of the landing

site determines the required orbital inclination by the relation thatInclination = Latitude + Offset. However, as was shown previously for a speci-fied transfer trajectory, the nodal position is determined once the orbital incli-nation is chosen. Hence, if this type of landing maneuver is utilized, thelanding-site location on any parallel of latitude is uniquely determined by thetransfer trajectory characteristics and the offset. Thus, if a maximum stay timeis to be obtained for fixed offset, a rather strong restriction on the possiblelanding sites must be accepted.

The landing-site capability can be increased at the expense of decreasingthe allowable stay time on the lunar surface by a slight modification of thelanding procedure illustrated in figure 4. Instead of landing at the most extremepoint to the left, the landing can be made at some point between the two extremesa l o n g t h e d e s i r e d p a r a l l e l o f l a t i t u d e . A g a i n , t h e e x p l o r a t i o n t i m e i s l i m i t e dto the time it takes for the landing site to move to the right-hand limit point.With this variation of the original landing procedure there is a range of possiblelanding sites along each parallel of latitude, and each landing site has a certainstay time associated with it.

The location of the possible landing sites on the lunar surface and the staytime at each site for any nominal earth-moon transfer trajectory can be calculatedusing equations (1), (2), and (3), together with the results given in figure 3.Figure 5 gives the landing-site stay-time restriction for a median-energy, low-inclination transfer trajectory. The figure shows the face of the moon as seenfrom the earth. The boundaries are the locus of points on the lunar surface atwhich landings can be made and at which the lunar excursion module can stay forthe specified time without requiring landing and take-off offsets of more than 50.The hatched line is the line of maximum stay time corresponding to the landingprocedure illustrated in figure 4. In the equatorial and polar regions theexcursion module can stay on the surface indefinitely and have the capability ofreturning to the command module within the limiting amount of offset; for example,if it is possible to establish an equatorial lunar orbit, then the excursionmodule can land within 5 0 of the lunar equator, and the rotation of the moon onits axis will not affect the relative position of the excursion module withrespect to the orbital plane of the command module.


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l

It is seen that a considerable portion of the lunar surface is subject tomanned exploration for periods of a few days or less. However, using this landingprocedure with the specified offset and the specified nominal transfer trajectorywill not allow landings in the midlatitudes on the western limb of the moon.Before the landing capability can be extended into this region, either the offset,the transfer trajectory characteristic, or the landing procedure must be changed.

Increasing the offset will not be considered since this parameter is directlyrelated to the propulsion capabilities of the lunar excursion module, and a siza-ble increase above the value used here is not expected. In addition, changingthe transfer trajectory characteristics will not yield a large increase in thepossible landing area because the location of the landing sites on the lunar sur-face is essentially determined by equation (1), and the results given in figure 3indicate that for manned missions, the location of the entry points is limited toa s m a l l r e g i o n o n t h e s p h e r e o f i n f l u e n c e . T h u s , t h e v a l u e s o f ^ a n d T 1 t h a tappear in equation (1) can only take on variations of about 8 0 from the meanvalues used to obtain the results of figure 5, and consequently changing thetransfer trajectory characteristics would result in a displacement of the bound-aries shown in figure 5 by only t 8 0 in both latitude and longitude.

Finally, there are a number of ways in which the landing procedure can bealtered so that the excursion module can land at the midlatitudes on the westernlimb of the moon. First, the landing procedure considered heretofore was designedso that the excursion module could return to the command module at any time duringt h e e x p l o r a t i o n p e r i o d w i t h o u t r e q u ir i n g a n o f f s e t o f m o r e t h a n 8 . A s m i s s i o nexperience is gained, such a requirement may be relaxed in order to increase thelanding-site capabilities inasmuch as without this requirement, it is possible toland at any point on the lunar surface. The motion of the moon may take theexcursion module a large distance out of the orbital plane of the command module;however, twice during the lunar month the orbital plane will pass over the landingsite and the return flight can be initiated. The exploration time at each pointon the surface is fixed by simple geometrical constraints and again by the resultsillustrated in figure 3 and equation (1).

Consider again landing procedures which allow return during the entireexploration period. Note that for a specified landing site on the western limbit is possible to establish a lunar orbit such that after landing, the excursionm o d u l e w i l l h a v e a d i s p l a c e m e n t o f 8 t o t h e w e s t o f t h e o r b i t a l p l a n e o f t h ecommand module; however, the inclination will be greater than the sum of the off-set and the latitude. The moon's rotation will again cause the landing site tomove eastward relative to the orbital plane; and after a short angular travel,t h e l a n d i n g s i t e w i l l b e e a s t o f t h e o r b i t a l p l a n e w i t h a n o f f s e t S a n d t h eoffset will be increasing. If the excursion module returns to the command moduleat this time, the exploration period in days is given by;

I' = 1 sin-sin ^ cos i + sin 8 _sin-1 s i n A cos i - sin 8)13.2os T sin ios A s i n i


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This equation gives an exploration time of at least 3.6 hours per degree of off-s e t . T h e r e f o r e , w i t h t h i s l a n d i n g p r o c e d u r e a n d a 5 0ffset, every point on thewestern limb can be explored for at least 18 hours.

Although it has not been stated explicitly, it has been assumed that thelunar excursion module initiates its landing maneuver soon after the lunar orbitis established. Suppose, however, that the excursion module does not land imme-d i a t e l y . B e c a u s e o f t h e m o o n ' s m o t i o n s , t h e n o d a l p o s i t i o n o f t h e o r b i t a l p l a n eof the lunar excursion module will appear to move westward along the lunar equa-tor. If after some specified waiting time in orbit the landing procedure illus-trated in figure 4 is initiated, the nodal position will have precessed westwardrelative to the lunar surface from its original position through an angle equalto the product of the moon's rotational rate and the waiting time in orbit.Therefore, for a specified waiting period, the landing sites and the correspondingexploration times could be obtained from figure 5 by simply rotating the areas inthe figure westward through the aforementioned angle. To increase the area ofpossible landing sites to the entire western limb would require waiting times inorbit of about 6 days.

CONCLUDING REMARKS

For lunar missions utilizing the lunar-orbit-rendezvous technique, thedetermination of the possible lunar-landing sites and the corresponding explora-tion times is seen to depend first on the geometrical properties of lunar orbitswhich can be established from typical earth-moon transfer trajectories and sec-ondly on the particular lunar-landing procedure utilized. For a landing proce-dure which affords mission abort during the entire exploration period, a largeregion of the moon can be explored for periods of a few days or less. A sizableincrease in the possible landing-site area on the lunar surface cannot be obtainedby changing the characteristics of the transfer trajectory; however, there are anumber of variations in the basic landing procedure which allow landings to bemade at any point on the lunar surface.

REFERENCES

1. Plummer, H. C.: An Introductory Treatise on Dynamical Astronomy. DoverPubl., Inc., 1960.

2. T o l s o n , R o b e r t H .: G e o m e t r i c a l C h a r a c t e r i s t i c s o f L u n a r O r b i t s E s t a b l i s h e dFrom Earth-Moon Trajectories. NASA TN D -1780, 1963.

6


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o"^*~^~^^ v^~^^^ESTABLISHMENT OF LUNAR ORBIT

MOON

'UNAR^~~~~LANDING SITESMOTION ~~_~~ TRANSFER TRAJECTORYt '-ARTHi=LUNAR-ORBIT INCLINATION TOTHE EARTH-MOON PLANED=4SCENDING NODAL POSITIONOF LUNAR ORBITI ,TRANSFER-TR4JECTORY INCLINA-TION TO THE EARTH-MOON PLANE

Figure II NJECTION

LAUNCIPOINT

POINT

GEOMETRICAL CONSTRAINTS ON LUNAR ORBITS

NINJECTION ENERG Y>c^ ' \FTRAJECTORY INCL I NAT I ON) ' ~~~X -`^~JNFLUENCE~~

ENTRY POINT \~TO

MOON'S

ARTHMOTIONFigure 2

7


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LOCATION OF ENTRY POINTS ON LUNAR SPHEREOF INFLUENCE

LUNAR SPHERE OF INFLUENCERADUS5 8 ,0 0 0 K M )MOONSMOTONSTOEARTHL A T I T U D E , T R A N S F E R - T R A J E C T O R Y7 7 , D E GN C L I N A T I O N , D E G1 0030N J E C T 1 0 N - V E L O C I T Y09 9 2, 9 9 4996 1.0A T I O3 0-1 0906 0LONGITUDE, , DEGFigure 3EXPLORATION TIME FROM RENDEZVOUS CONSIDERATIONS

EXPLORATION TIME=TAN 7GO DAYSNi =SNMC3 . 2A P A R A L L E L O F\= + sL A T I T U D ETHROUGHLANDING SITEs

1_UNARQ U A T O RTO EARTHL U NA R SU R FA CE Figure 48


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L A N D I N G S I T E -S T A Y T I M E R E S T R I C T IO N S F R O MR E N D E Z V O U S C O N S I D E R A T IO N S , 8 = 5 0N

W E

Figure

9


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2. RETURN-TO-EARTH CONSIDERATIONS AND STAY-TIME CONSTRAINTSFOR THE LUNAR MISSIONBy John P. Gapcynski

SUMMARY

This paper presents a discussion of the restrictions which are imposed on theachieve a satisfactory return flight to a

earth-moon transfer tra-return flight is discussed.

INTRODUCTION

The return flight from the moon is one of the more important phases of the

(defined here as the reentry range).If it is assumed that the vehicle is in a circular orbit about the moon at

by the two parametersinclination of this orbit to the earth-moon

earth-return tra-injection velocity and total mission

EARTH REENTRY CONSTRAINTS

In solving the return problem, it is necessary to work back from the earthrequirements for earth reentry and then

vehicle must achieve

1 1


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- - - - - - - - - - - -constraints which causes the most trouble because the specification of a desiredtouchdown point introduces the factor of time into the problem.

In satisfying the requirement of a specific touchdown longitude, the -param-eters of importance are the time selected for injection and the return-trajectoryflight time. In satisfying the requirement of a specific touchdown latitude, the-parameter of importance is the inclination of the return-trajectory -plane to thee a r t h - m o o n p l a n e . T h i s p a r a m e t e r i s d e s i g n a t e d b y I i n f i g u r e 1 . N o t e t h a tthis parameter I is the inclination of the return trajectory to the earth-moonplane and should not be confused with the inclination of the lunar orbit to thee a r t h - m o o n p l a n e d e s i g n a te d i n f i g u r e 1 b y i . I t i s w e l l t o n o t e h e r e t h a t i tis possible to obtain one specific value of the return-trajectory inclinationfrom lunar orbits which have a wide range of values of orbital inclination tothe earth-moon plane.

To illustrate the importance of the return-trajectory inclination, considert h e d a t a s h o w n i n f i g u r e 2 . I n t h i s f i g u r e i s s h o w n t h e v a r i a t i o n o f t h e r e q u i r e dreturn-trajectory inclination I with the angular position of the moon in its orbitto achieve a touchdown latitude of 3 00. Th e a b s c i s s a i s e s s e n t i a l l y a t i m e s c a l e;that is . ,uring the lunar month, the angular position of the moon will changet h r o u g h 3600n t h e d i r e c t i o n s h o w n a t t h e r a t e of 1 3 .20 p e r d a y . M a x i m u m p o s i -tive, or northerly, declination of the moon on this scale occurs at an angularp o s i t i o n o f 900., and maximum negative, or southerly, declination occurs at anangular position of 2700. Results are presented for representative values ofthe reentry range of 3 00 ,0 O , nd 900.

Note from these results that for any particular reentry range the return-trajectory inclination which is required changes rapidly with the position of them o o n . F o r e x a m p l e ., c o n s i d e r t h e r e sul t s fo r a 30 0 r e e n t r y r a n g e . I f t h e r e t u r nflight is initiated when the moon has an angular position of about 230 0 , then thereturn trajectory must be inclined 70 0 t o t h e e a r t h - m o o n p l a n e . T h r e e d a y s l a t e r ,when the moon has a position of about 270 0 , the return trajectory must be inclinedo n l y 10 0 to the earth-moon plane.

Further, the time period during the month when a return flight to this lati-t u d e i s p o s s i b l e i s a f u n c t i o n o f t h e r e e n t r y r a n g e a v a i l a b l e . F o r e x a m p l e , w i t ha 300 reentry range a return flight is possible only when the moon is near itsp o s i t i o n o f m a x i m u m n e g a t i v e d e c l i n a t i o n . W i t h a 6 0 0 reentry range approximatelyhalf a month is available ., and with a 90 reentry range a return flight can bemade at any time during the month.

These time intervals place no constraints on the magnitude of the returni n c l i n a t i o n r e q u i r e d . I t i s d e s i r a b l e , however, to utilize return trajectorieswith low inclinations to the earth-moon plane, not only from a guidance stand-point, but also because of the fact that return flights initiated from lunarorbits having low orbital inclinations may require low return-trajectory inclina-t i o n s . I n o r d e r t h a t a t o u c h d o w n l a t i t u d e o f 300e achieved with reasonablevalues of the return inclination, the mission should be initiated when the moonis approaching its position of maximum negative declination. This time restric-tion may be alleviated somewhat if an additional touchdown point is assignedwhich has a southern latitude of 3 00. The variations in this case are similar1 2

IT


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'^..01111.,rto those shown in figure 2, except for a shift along the abscissa of 1 80 0 , andreasonable values of the return inclination may now be obtained when the moonis approaching its position of maximum positive declination.

In summary, the return trajectory must be such that the vehicle will reenterthe earth's atmosphere within a desired corridor, and it must be inclined to theearth-moon plane at some specific angle which depends upon the time of the monthselected for the initiation of the return flight and the reentry range available.The problem of the return mission, then, is one of determining how these require-ments may be met or, alternately, of determining the values of the return-trajectory inclination which can be achieved from an arbitrarily oriented lunarorbit

ANALYSIS

It is convenient in a parametric study of this type to utilize the "patched-conic" type of solution (ref. 1), wherein the return vehicle is assumed to besubject only to the attraction of the moon within the lunar sphere of influenceand subject only to the attraction of the earth exterior to this region. Thus,the problem becomes a combination of two solutions, one geocentric and the otherselenocentric. With this type of approach, it is possible to determine the con-ditions at the lunar sphere of influence which will satisfy the reentry con-straints. When these conditions are known, it is then possible to determine thelunar orbital characteristics which satisfy them. Further, it will be assumed inthe discussion which follows that the vehicle exits normal to the lunar sphere ofinfluence. Actually, the angle between the vehicle velocity vector and radiusvector at this point is of the order of 3.5 0 , a n d r e sul t s o b t a i n e d fr o m t h i sstudy, without the use of this assumption, are presented in reference 2. How-ever, by using this assumption here it is possible to obtain a very good phys-ical picture of the phenomena associated with the return flight,

RESULTS AND DISCUSSION

Sphere of Influence ConditionsThe types of results which may be obtained from the analysis of this problem

are shown in figure 3. For orientation purposes, a sketch of the lunar sphere ofinfluence is shown in the upper portion of the figure. The direction to earthand the direction of the moon's motion are indicated by arrows. Two angles whichwill be referred to are a, or the longitude of a position on the lunar sphere ofinfluence, measured from the earth-moon line, and 7 1 or the latitude of a posi-tion on the lunar sphere of influence, measured from the earth-moon plane.

In the lower portion of figure 3 are shown the loci of vehicle exit pointson the sphere of influence which result in a satisfactory earth return. Theordinate is the latitude of the exit position, and the abscissa is the longitudeof the exit position. Only representative values of the return inclination of


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d51 0 0 , 4o O , nd 90 0r e s h o w n . T h e l o c i f o r a l l o t h e r r e t u r n i n c l i n a t i o n s w i l lfall between the 90 0ariations.

If the vehicle exits from the sphere of influence anywhere along one of theseloci, then the return trajectory will satisfy the reentry-corridor constraint andhave the inclination specified. If the exit point has a negative latitude, thevehicle will return to the earth's Northern Hemisphere. If the exit point has apositive latitude, the vehicle will return to the earth's Southern Hemisphere, andthe return trajectory will be oriented as shown by the dashed line in figure 1 .

The implication of the results shown in figure 3 is that in order to returnto the earth the vehicle must exit from the trailing edge of the lunar sphere ofinfluence within a rather narrow band defined by latitude limits of approxi-mately 10 0 . For orientation purposes, this area is indicated by the shadedregion in the upper portion of the figure.

The parameter which distinguishes one exit point from another along any oneof these curves is the vehicle injection velocity increment required to achieveearth return. The minimum velocity increment is associated with exit positionsnear 80 0f longitude. As the exit position is shifted from this point in eitherdirection, the required velocity increment increases. The total mission flighttime - that is, the time from injection to reentry - decreases as the exit posi-tion is shifted toward the earth-moon line and increases as the exit point isshifted in the other direction. The reason for this increase in flight time forexit points having longitude values greater than 80 0s that the vehicle isheading away from the earth when it leaves the sphere of influence.

In view of the fact that there will be injection-velocity and flight-timelimits imposed upon the return mission, the actual exit area will be smaller thanthat discussed so far. If the arbitrary, but representative, limits of 3, 1 00 feetper second in the velocity increment and 1 00 hours in flight time are imposed, theexit area will be confined to the shaded region shown superimposed on the resultsin the lower portion of figure 3. The left boundary of this region is defined bythe velocity limit of 3, 1 00 feet per second and the right boundary, by the timelimit of 1 00 hours. An increase in the allowable value of either of these twoparameters will increase the size of this shaded area.

For a successful return mission within the defined limits of velocity andflight time, then, the selenocentric, or lunar trajectory, characteristics mustbe such that the vehicle exits from the sphere of influence within this shadedregion. The exact location of the exit point within this region will be deter-mined by the return-trajectory inclination desired and the velocity required toa c h i eve a s p e c i fi c l o ngi tud e .

Lunar-Orbit CharacteristicsThe required lunar-orbit characteristics may be assessed from an examinationof the results shown in figure 4. he shaded region in this figure is the per-missible exit area as defined previously. Superimposed on this exit area are the

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,.ateto0 , 20 0 , and 6.50.

The point to be made here is that since there are any number of arbitrarilyit is possible to return

cannot be less than the latitude ofarbitrarily the nodal

value that the orbital planerequired exit area.For example, if the desired inclination of the orbit is 20 0 , a return flightnodal positions having longitudinal values of approximately 800.

for the dashed lineapproximately 120 0 ), the trace of this orbital plane does not intersect theinclination at this

s u c h i s t h e c a s e: f i r s t ,node to the proper position, and4 will appear to precessof about 1 3.2 0 per day, and, therefore, there willtrace of the orbital plane

One additional point may be noted with relation to this apparent precession

time period that the trace of the orbit inter-

Further, it is to be noted that it is not possible to achieve exit pointslunar orbit has a low inclination to

w i t h l a t i t u d e s g r e a t e r t h a n 6.50inclined 6.50o the earth-moonthat the maximum value of the return-trajectory inclina-

period during the month whenp o s s i b l e .In summary, it has been shown that it is possible to achieve a satisfactory

once the inclination of the orbit has beenregion to

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As outlined by Robert H. Tolson in paper no. 1 , the use of the lunar-orbit-rendezvous technique to achieve landing sites anywhere on the lunar surface neces-sitates the establishment of lunar orbits having a range of inclinations from0 0 t o 90 0, nd associated with each inclination in this range is a specific nodalposition which is determined by the earth-moon transfer-trajectory characteris-tics. Similarly, it has been shown in this paper that return flights may beachieved from lunar orbits having inclinations in the range from 0 0 t o 90 0, ndthat here again a specific nodal position is required for each inclination,depending in this case on the velocity and time selected for injection into thereturn trajectory.

The problem remaining is to determine the relationship between the orienta-tion of the orbit when it is first established and the orientation required forthe return flight. This information is of fundamental importance to the lunar-orbit-rendezvous technique since the selection of lunar landing sites and theassociated stay times should be compatible with both the orbit establishment phaseand the return phase of the mission.

To illustrate this relationship, consider the results shown in figure 5.First, for orientation purposes, consider the sketch of the moon, with a vehiclein some arbitrary circular orbit, shown in the upper right-hand portion of thefigure. The darkened regions in this sketch represent the required range ofnodal-line positions in the earth-moon plane for lunar orbits having inclinationsb e t w e e n 2 0 0 a n d 90 0. f this area is viewed from the northern lunar latitudes,the result shown in the major portion of figure 5 will be obtained, where thedirection of the moon's motion and the earth are again denoted by arrows.

T h e h a t c h e d a r e a l y i n g b e t w e e n 1 3 5 0 and 1+5 0 represents the nodal-line posi-tions of those lunar orbits established from a typical earth-moon transfer tra-jectory and having inclinations between 20 0 and 90 0 . T h e s t i p p l e d a r e a l y i n gbetween 30 0 a n d 1 0 0 0 represents the range of nodal-line positions required for areturn flight to either the Northern or Southern Hemisphere of the earth, withinthe velocity and flight-time limits discussed previously and for the same incli-nation range, that is, between 20 0 and 900.

It is evident that the two areas are not compatible, and, therefore, areturn flight cannot be initiated immediately from any lunar orbit having aninclination greater than 20 0 , unless flight times of well over 1 00 hours areacceptable. Because of the motion of the moon, the nodal line of any orbit estab-lished in the hatched area will appear to shift in the direction shown at the rateof about 1 3.2 0 per day, and, therefore, after some specified waiting period thenodal line will have the correct orientation so that a return flight is possible.Note that this may require a waiting period from 3 to 4 days.

As the orbital inclination is decreased below 20 0 , both regions begin toexpand, and below 1 0 0 they overlap; thus, a return flight may be ini tiated at anytime after orbit establishment subject to the capability of obtaining the desiredreturn-trajectory inclination and achieving a specified longitude.

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^" m o

THE LUNAR RETURN PROBLEM

MOONMOON'SMOTION

RETURN TRAJECTORYEARTHREENTRY I = LUNAR-ORBIT INCLINATION

RANGEO THE EARTH-MOONPLANE

~^ R ^=ASCEND|NG NODAL POSITION-F LUNAR ORBIT----~RETURN -TRAJECTORY INCLINA-TION TO THE EARTH-MOONPLANEFigure IREQUIREDRETURN-TRAJECTORY |NCL|N4TK}N

TOUCHDOW N LATITUDE = 30 0

RETURN-TRAJECTORYOON'S MOTIONINCLINATION, I,3i20/DAYDE G120 r---->

80-EENTRY RANGE, DEG0-ECLINATIONAX. NEG.0020160 200 240 280 320 360ANGULAR POSITION OF MOON, DEG DECLINATION

Figure 2

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n * w *** * ~ ^ o ~ o w= " ~ e , ~ ~^~ s^, ^* * ^^ ^ ^^*, "U]C| OF EXIT POINTS ON THE LUNAR SPHEREOF|NFLUENCEMOON'SUNARMOTIONPHERE OFINFLUENCETO EARTHj^VELOCITY INCREASESELOCITY INCREASESTIME DECREASES-----TIMEINCREASES20-

LATITUDE, 77,D E G

-2O^-9v00 40 60 80 100 120 140 160 180LONGITUDE, u,DEGFigure 3LUNAR-ORBIT CHARACTERISTICS FOR THE RETURN W4|SS|DN

DEG92O20.5LATITUDE 77,ODEG

LONGITUDE, a , DEG

Figure 4

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MO O N SMOTION- A P P A R E N To NODE SHIFT1 0 03.2- /DAYORIENTATION OF LUNAR ORBITS20 < i : 5 90 0N O D A L - L I N E O R I E N T A T I O NF O RO R O R B ITR E T U R N - - \ E S T A B L I S H M E N T , M O O N ' SM O T I O NT O145* EARTH1 3 5 03 0 0

T O

Figure 5

- - - - - - - - - - - - - - - - - - - - - -20


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a Fr

3 A SIMPLE GUIDANCE TECHNIQUE FOR LUNAR LETDOWNBy Richard Reid

SUMMARY

The back-up system proposed herein would allow the pilot of the lunar excur-orbit transfer,

where the braking descent was tosimple thrust maneuver and attain the preselected

t h e b r a k i ng d e s c e n t .

INTRODUCTION

Studies are underway at the Langley Research Center to determine methods of

wherein the pilot of the lunar excursion module (LEM) mon-pilot to monitor

assure arrival at a preselected

GENERAL CONSIDERATIONS AND PROCEDURES

In this study it is assumed that the Apollo vehicle has established an7-nautical-mile circular orbit. At a preselected point a thrust maneuver isaltitude at pericynthion94.4.n this study the range angle is

The magnitude of the error in thrust is the only error parameter consideredused in this

944, here braking to the lunarpilot from an optical

vehicle in its known orbitradar altimeter or

d e t e r m i n e w h a t9440.f his altitude will not be

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f k ^.sufficiently near 50,000 feet at 94.4 Y a correction is made by applying thrustalong the local vertical. The procedure of making a midcourse correction stand-ardizes the lunar-landing maneuver. Since the corrections required will be small,thrust along the local vertical is used for simplicity. This thrust can beapplied with the reaction control motors.

RESULTS

T h i s m i s s i o n p r o f i l e i s s h o w n i n f i g u r e 1 . T h e o p t i m u m m a n e u v e r f o r t h eLEM orbit transfer is shown by the dashed line. If too much thrust is applied,the LEM enters an "error orbit" which results in an altitude of less than5 0 , 0 0 0 f e e t a t 0 = 9 4 . 4 0 u n l e s s a c o r r e c t i o n i s m a d e . T h e f i g u r e s h o w s a c o r -r e c t i o n a t 0 = 60 0 t o p l a c e t h e L E M a t 5 0 , 0 0 0 f e e t a t 0 = 9 4 . 4 0 . If such amidcourse correction is made, the LEM is no longer in a synchronous orbit. Sub-sequent corrections will return the LEM to the original synchronous orbit if thepilot decides not to land.

F i g u r e 2 s h o w s t h e e r r o r s i n a l t i t u d e a t 0 = 9 4 . 4 0 t h a t w o u l d r e s u l t f r o mthrust errors (expressed as errors in velocity change LV) in the orbit-transferm a n e u v e r . Z e r o e r r o r r e s u l t s i n a n a l t i t u d e o f 5 0 , 0 0 0 f e e t a t 9 4 . 4 0 . An errorof approximately 50 fps in the transfer maneuver would result in the LEM hittingthe lunar surface at 94.40.

Figures 3(a), 3(b), and 3(c) are examples of the charts a pilot would use tod e t e r m i n e w h a t h i s a l t i t u d e w o u l d b e a t 0 = 9 4 . 4 0 b y t a k i n g a l t i t u d e m e a s u r e -ments at range angles of 3 0 0 , 4 5 0 , and 60 0 , r e s p e c t i v e l y . I n a d d i t i o n , t h e s ec h a r t s s h o w t h e D U c o r r e c t i o n s r e q u i r e d t o r e a c h 5 0 , 0 0 0 f e e t a t 9 4 . 4 0 , by meansof vertical thrust, if the pilot determines that there is an error in the orbitt r a n s f e r . T h e h a t c h e d b a n d s i n t h e f i g u r e s i n d i c a t e a 1 p e r c e n t i n a c c u r a c y i naltitude determination and the consequent inaccuracies in altitude at 0 = 94.40a n d i n c o r r e c t i o n - v e l o c i t y r e q u i r e m e n t s . B y c o m p a r i n g f i g u r e s 3( a), 3(b), a n d3(c) it is seen that correction requirements are smaller for the smaller rangeangles, but inaccuracies in altitude determination are greater for the smallerrange angles.

CONCLUDING REMARKS

It is seen that an error in the orbit-transfer maneuver does not necessarilycreate an abort situation. By monitoring the coast phase, the pilot can detecterrors, establish correction-velocity requirements, and perform a simple thrustmaneuver to reach preselected conditions for the start of the braking descent.Corrections can be made that do not jeopardize the mission, since subsequent cor-rections can return the LEM to the original synchronous orbit.

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3,A "A it.a s aREFERENCES

. Queijo, M. J., and Riley, Donald R.: A Fixed-Base-Simulator Study of theAbility of a Pilot to Establish Close Orbits Around the Moon. NASA TN D -917 ,1961.Houbolt, John C., Bird, John D., and Queijo, Manuel J.: Guidance and Naviga-tion Aspects of Space Rendezvous. Proceedings of the NASA-University Con-

ference on the Science and Technology of Space Exploration, Vol. 1 , NASAS P - 1 1 , 1962, pp. 353-366. (Also available as NASA SP-17.)


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MISSION PROFILEORBIT TRANSFER

\`'

^SYNCHRONOUS ORBIT

A P O L L O O R B I TE R R O R O R B I T5 0 , 0 0 0 F T

Figure 1

EFFECTS OF VELOCITY ERROR

30AVE R R O R ,FPS

1 5

O------- --I \-150 10 20 30 40 50 60 70 X 103A L T I T U DE A T 6 = 9 4 .4 , F T

Figure 2

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P R E D IC T IO N O F A L T IT U D E A N D C O R R E C T I O N V E L O C I T YALTITUDE AT 8= 30 0 , FT305 x103

300 NOMINAL295RANSFER290285 -20-100 CORRECTION,1020280

A l 30275 iI270I2650 10 20 30 40 50 60 70 X 103ALTITUDE AT 8 = 94.4 0 , FTFigure 3(a)P R E D I C T IO N O F A L T IT U D E A N D C O R R E C T I O N V E L O C I TYALTITUDE AT 8=45, FT220 X103

210 -25NOMINAL -12.5200:.---TRANSFER_ _ _ _ OV1% CORRECTION,190 12.5PS25180 i

40

1701600 i 70 X103000 40 50 60ALTITUDE AT 8=94 .4, FTFigure 3 (b)

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0VC O R R E C T I O N ,F P S3060

12 5

110

95

P R E D IC T IO N O F A L T IT U D E A N D C O R R E C T IO N V E L O C I TYALTITUDE AT 8 = 60*, FT

140, x 10 330800 10 20 3040 50 607 0 x 1 0 3

ALTITUDE AT 8=94.4*, FT

Figure3(c)

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LUNAR LANDING


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4. VISUAL SIMULATION STUDY OF LUNAR HOVERING,TRANSLATION, AND TOUCHDOWN

By Maxwell W. Goode

SUMMARY

A pilot's ability to control hovering, translation, and touchdown on the

a side-arm controller.The results of this investigation indicate that the hovering, translation,capability of a human

inherent in V/STOL vehicles operating in the earth gravity field.

0.5 radian/sec 2 seems adequate for a rate com-3 feet/sec 2 n e a r t h e

INTRODUCTION

This investigation was a preliminary study of a pilot's ability to control

of the touchdown point will be made by

It was assumed for this study that the landing vehicle had been brought todeorbit

DESCRIPTION OF APPARATUS

The simulation used in this study was a three-degree-of-freedom system ori-

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in the vertical plane with a throttle and a side-arm controller. The throttlecontrolled the rate of fuel consumption of a fixed rocket engine with a maximuminitial thrust-weight ratio of 0.4g. The attitude was controlled by a propor-tional rate command system.

Figure 1 is a picture of the simulator showing the throttle, the side-armc o n t r o l l e r , a n d t h e d i s p l a y p r e s e n t e d t o t h e p i l o t . A l t i t u d e w a s d i s p l a y e d a l o n gthe Y-axis and horizontal distance along the X-axis. Vehicle position and atti-tude were indicated by the meter in the upper right corner of the display, shownas a closeup in figure 2. The position indicator was a pointer only and did notm a k e a v i s i b l e t r a c e o f t h e r u n s . V e h i c l e a t t i t u d e i n t h e v e r t i c a l p l a n e w a sindicated by deflection of the needle from the vertical. A fuel-quantity indi-cator was also included.

This simulation was oriented toward contact flight in that the pilot had tojudge his acceleration, velocities, and displacements from the motions of thed i s p l a y . T h e b a c k g r o u n d o f t h e d i s p l a y w a s b l a n k , w i t h o n l y t h e s u r f a c e a n ddesired landing points included. The success of a landing was determined by thetouchdown conditions. The established limits f or touchdown were: horizontalvelocity, less than 5 ft/sec; vertical velocity, less than 1 0 ft/sec; touchdownangle, less than 15 0 ; a n d m i s s d i s t a n c e, l e s s t h a n 40 fe e t .

RESULTS

F i g u r e 3 is a graphic presentation of several landing trajectories from a1 , 0 0 0 - f o o t i n i t i a l a l t i t u d e a n d a 1 , 0 0 0 - f o o t t r a n s l a t i o n d i s t a n c e . A n i n i t i a laltitude of 500 f e e t a n d a t r a n s l a t i o n d i s t a n c e o f 6 0 0 f e e t w e r e a l s o u s e d . A nanalytical study (reported in ref. 1 ) indicated that a ballistic-type trajectorywas approximately 15 to 20 percent more economical than trajectories requiringcontinuous thrust. The fuel expenditure for the entire maneuver, however, repre-sents only 2 to 3 percent of the initial hovering weight of the vehicle.

Trace 0 s a b a l l i s t i c t r aj e c t o r y . T o f l y t h i s t r aj e c t o r y t h e p i l o tpitched the vehicle forward and applied maximum thrust for a short time. Heyt h e n s h u t o f f t h e e n g i n e a n d c o a s t e d . A s t h e v e h i c l e a p p r o a c h e d t h e s u r f a c e i nthe vicinity of the landing point, the pilot pitched the vehicle back, appliedm a x i m u m t h r u s t t o b r a k e , a n d m a d e t h e l a n d i n g . T h i s t r aj e c t o r y h a d t w o d i s t i n c tdisadvantages: (1) There were too many parameters, such as th e pitch angle, thefiring time, and the correct braking maneuver, that had to be judged correctlyin order to capitalize on any fuel economy, and (2) the vehicle was not underpositive control at all times.

T r a c e O 2 i s a m o d i f i e d b a l l i s t i c t r a je c t o r y . I n t h i s o n e t h e p i l o t p i t c h e dthe vehicle forward, and after a few seconds he applied maximum thrust and thenshut off the engine. In the touchdown region he pitched the vehicle back andapplied maximum thrust again to brake and make the landing. This trajectory hadthe same disadvantages as the true ballistic trajectory.

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T r a c e i s a h e l i c o p t e r - t y p e a p p r o a c h . T h e p i l o t r e d u c e d p o w e r c o n s i d e r -vertical attitude and started arresting his

the horizontal velocity and made the

to correct any miss distance before thevertical attitude before touchdown.

T r a c e O s the trajectory that showed the most promise and is the one that

(a) Pitch the vehicle forward approximately 30 0o 350.(b) Reduce power slightly to fly a near line-of-sight path to a point esti-

touchdown point.(c) As the halfway point is reached, pitch the vehicle back approximately00o 35 0o start braking.(d) Manipulate the attitude of the vehicle when approaching the touchdown

(e) Add sufficient power to arrest the vertical velocity and make the

This trajectory was under positive control at all times and could be flownor vehicle

A series of preliminary runs was made with trajectory Q o determine anradian/sec 2 and a maximumr a d i a n /s e c . F i g u r e 4 is a curve of the attitude-control

1 ) T h e p r e d o m i n a n t b r a k i n g p o i n t a t 0.5 radian/sec 2 or less, used inpercent of the runs, and (2) the infrequent use of maximum available power.maximum power was used was in the turn-around

0.5 radian/sec 2 and maximum angular velocity of 1 radian/sec.For fuel with a specific impulse of 315 s e c o n d s, a fue l w e igh t o f a p p r o x i -3.3 percent of the initial hovering weight of the landing craft allowed the

required for translation and touchdown

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xx . : xwas approximately 1 minute, and the fuel consumed was about the same as for anequivalent period of hovering.F i g u r e 5 shows the distribution of fuel usage as a function of translationdistance for two typical runs. Approximately the same total fuel was requiredfor the two runs, and the curves indicate that the distribution of fuel usagew a s a b o u t t h e s a m e . S t a r t i n g a n d b r a k i n g t o a l a n d i n g c o n s u m e d a l a r g e p o r t i o nof the fuel - 15 to 20 percent for starting and 35 to 45 percent for braking andlanding. Thus a relatively large change in translation distance would requireonly a small change in total fuel.

One of the most significant results of this study was the disclosure of acontrol coupling between vehicle attitude and thrust. Generally speaking, atti-tude controlled the translational conditions, and thrust controlled the altitudeconditions; but the relatively large pitch angles involved caused pronouncedcross coupling. To help explain the source of the coupling problem, figure 6 isan illustration of the force vectors of a rocket-propelled landing vehicle. Thethrust for hovering on the moon will be 1/6 of that required on earth because oft he 1/6 gravity term. To translate with a fixed rocket-engine system in theearth gravity field, a pitch angle of 5.70s required to produce a horizontalacceleration of l/lOg with no appreciable change in thrust to maintain altitude.However, in the lunar gravity field a pitch angle of 31 0ill be required with anincrease of 17 p e r c e n t i n t h r u s t t o m a i n t a i n a l t i t u d e . A t t h i s p i t c h a n g l ea p p r o x i m a t e l y 50 percent of any thrust change is realized in horizontal accelera-t i o n a s c o m p a r e d w i t h a p p r o x i m a t e l y 1 0 p e r c e n t f o r t h e e a r t h c a s e . T h i s c o u p l i n gmanifested itself in ballooning or overshooting the landing point. It was, gen-erally, better to choose another suitable landing point rather than to try tocorrect an overshoot by backing up, especially when very close to the surface.The coupling was not necessarily uncontrollable, but suggests a restriction ofp i t c h a ngl e c l o s e t o t h e sur fa c e .

F i g u r e 7shows the distribution of maximum horizontal acceleration below analtitude of 1 00 feet. In most of the runs the maximum horizontal accelerationw a s 2 t o 3 ft/sec 2 o r l e s s . H o r iz o n t a l a c c e l e r a t i o n w a s a c h i e v e d b y t i l t i n g t h evehicle in this study; however, if horizontal-firing engines were considered asa solution to the coupling problem, engines of the 1 ,000-pound-thrust class wouldbe required to produce an acceleration of about 3 ft/sec 2 for a vehicle weighinga b out 10,000 e a r t h p oun d s .

CONCL USIONSThis study of piloted control of hovering, translation, and touchdown on thelunar surface has yielded the following conclusions;1. Pilots preferred to fly a line-of-sight type of maneuver from hovering tothe desired landing point.2. The effects of a control coupling between attitude and thrust can caused i ffi cul t y a n d p e r h a p s w a r r a n t fur t h e r s tud y .

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3- If the landing point is overshot, a correction by backing up should be4 . An angular acceleration of 0.5 radian/sec t seems adequate for attitude

5 . A horizontal-acceleration capability of 2 to 3 feet/sec t close to the

REFERENCE

,, Carl E., and Hamza, Vladimir: Analysis of CloseL u n a r T r a n s l a t i o n T e c h n i q u e s . N A S A T E R - 1 2 6 , 1 9 6 2 .

31. . . . . . . . . . . .


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S I M U L A T O R

Figure 1-2034-1ATTITUDE AND POSITION INDICATOR

Figure 2-2034-232


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1 , 0 0 0

A L T I T U D E , F T

(D

8 0

6 0

40

2 0

LANDING TRAJECTOR IES

1 , 0 0 0T R A N S L A T I O N D I S T A N C E , F T

Figure 3

ATTITUDE-CONTROL POWER DISTRIBUTIONP E R C E N T A G E O F R U N S A T O R B E L O W

100 r --102468.0M A X I M U M A N G U L A R A C C E L E R A T I O N U S E D , V , R A D I A N S / S E C 2Figure 4

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lo x 1 0 ,T R A N S L A T I O N D I S T A N C E , F TP E R C E N T F U E Llu u8 06 0402 00 FUEL USAGEFigure 5F O R C E V E C T O R S F O R R O C K E T - P R O P E L L E D L A N D I N G V E H I C L E

E A R T H ( g )U N A R ( g /6 )H O V E R I N GTHRUST, FF16IWEIGNT,W/6-1 t- 53*3 1 0TRANSLATIONTO MO VE 100 FEET/6IN ABOUT 12 SECSqto/loW g = 32.2 FT/SEC2gue-2o 34 -6

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SYMBOLS

AreaaccelerationForcegcceleration due to gravityIoment of inertiaZe ngt hmasstimeVelocityangular accelerationRr avi t a t i o n al r a t i oAeometric model scale factorot r e s s o f y i e l d s t r a pwngular velocity MODEL AND APPARATUSThe 1/6-scale dynamic model used for the investigation is shown in figure 1 .Simulated full-scale spacecraft height was 185inches, basic landing-gear diam-eter was 222 inches, and initial center-of-gravity height above groundline was108inches The basic landing gear was a symmetric four-point arrangement, asshown in figures 1 and 2(a). Each of the four legs consi sted of three strutsm o u n t e d s o a s t o f o r m a n i n v e r t e d t r i p o d . T h e u p p e r s t r u t t e l e s c o p e d d u r i n gimpact, and the lower V-strut was a hinged unit which served to guide and stabi-lize the tripod. Impact energy was absorbed by yielding a pure nickel strap withthe telescoping strut. The landing pads were ball-joint mounted. Verticalworking stroke for the full-scale gear was approximately 2 feet and was designedto give a maximum load of 2 earth g units at an impact speed of 15ft/sec.

Modifications to the basic landing-gear arrangement are illustrated in fig-ure 2. Figure 2(b) shows the asymmetric four-point arrangement in which the for-ward pads were extended in the direction of horizontal flight The radial dis-t a n c e t o t h e c e n t e r o f t h e o f f s e t p a d s w a s 33percent greater than that for the38


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2

i nof the pad was the same as for the basic

The scale relationships pertinent to the earth model tested are shown in1/6-scale model was used inasmuch as this was suitable

1/6- s c a l e m o d e l a n d t h e f ul l - s c a l ethus exact structural scaling ofr f o r c e s w a s p r o v i d e d . T h e g r a v i t a t i o n a l r a t i o R is dictated byearth's gravity is six times that of the moon,resulting

a dynamically scaled model.The investigation was conducted by launching the model as a free body with3 esired landing attitudes were

by adjusting the3.

The landing surfaces included a flat hard surface, a ledged hard surface, a

f 400 m i c r o n s . E s t i m a t e d b e a r i n g s t r e n g t h o f t h e l o o s e l y c o m p a c t e d p u m i c e40 lb/cu ft.lunar surface

RESULTS AND DISCUSSION

A l l d a t a p r e s e n t e d a r e f u l l - s c a l e v a l u e s . I m p a c t d a t a p r e s e n t e d a r e c o n -

Landing ImpactTypical accelerations measured at the center of gravity of the module with

figure 4.plotted( - 15 0nd 0 0 ), as. V e r t i c a l s p e e d f o r t h e s e d a t a i s c o n s t a n t a t 1 0 f t / se c . R e s u l t a n t0 to 33 0. he sliding coefficient of friction0.4. R e s u l t s f o r a p i t c h a t t i t u d e o f 15 0n d fo r

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landings astride the ledge were about the same as for those at a pitch attitudeo f - 15 0 . As expected, for a constant-force shock absorber, there was littlechange in maximum acceleration values over the entire speed range. Accelerationsfor all three configurations during hard-surface landings were within the limitsshown on this figure, and no overturning occurred.

Typical accelerations obtained during landings on a flat surface with pumice-dust overlay are shown in figure 5. Maximum normal, longitudinal, and angularaccelerations are plotted against variation of horizontal speed for pitch atti-tud e s o f 15 0 a n d f o r d u s t 5 a n d 2 1 i n c h e s d e e p . A g a i n , v e r t i c a l s p e e d i s c o n -s t a n t a t 1 0 f t/ s e c . T h e s e l a n d i n g s w e r e m a d e w i t h t w o l e g s f o r w a r d . A c c e l e r a -tion characteristics were generally similar to those for hard-surface landings.Negative longitudinal acceleration (or drag) increased by a factor of 2 or 3 toa b out 17g units with increase in horizontal speed. It was estimated that thecoefficient of friction was 0.7 to 1 .0 in pumice. Drag forces also decreasedrighting moments, as indicated by the reduced angular accelerations at a pitchattitude of -1 5 0 . These factors resulted in the overturning of the module in thedirection of the flight path, as indicated by the solid symbols. The sample datain figures 3 and 4 show that the maximum landing accelerations were small and didnot vary a great deal over a wide range of landing conditions. Maximum normalacceleration was 2g units, maximum longitudinal acceleration was 1 , g u nits, andmaximum angular acceleration was 1 2 2 radians/sec2.

Landing StabilityA comparison of overturn characteristics of the three configurations is pro-vided in figure 6 for various pitch attitudes and h orizontal velocities. These

characteristics were obtained for simulated landings in a 2 1 -inch pumice overlayat a constant vertical speed of 1 0 ft/sec. These landings were made with twolegs forward. Conditions where overturning occurred are shown by the solid sym-b o l s . T h e s y m m e t r i c a l f o u r - p o i n t c o n f i g u r a t i o n o v e r t u r n e d a t a h o r iz o n t a l s p e e dof 1 0 ft/sec or greater, depending upon touchdown attitude. The symmetrical five-point configuration increased overturn stability slightly, as can be seen by com-paring the circular and square data symbols at a pitch attitude of -1 5 0 a n d ah o r iz o n t a l s p e e d o f 1 0 f t/ s e c . H e r e t h e f o u r - p o i n t c o n f i g u r a t i o n o v e r t u r n e d a n dthe five-point configuration remained upright. With the asymmetric four-pointgear, there is a significant improvement in overturn stability, as can be seen bycomparing open triangles with the solid circles and squares at a horizontal speedof 1 5 ft/sec and at pitch attitudes from -1 5 0 t o 1 5 0 . T h e s y m m e t r i c f o u r - a n dfive-point configurations overturned, whereas the asymmetric configurationremained upright.

A summary of stability and motion characteristics illustrated in motion pic-tures of several typical test runs is presented in table II. During hard.-surfacelandings with horizontal speed, a more steady platform was obtained during andafter slideout by landing with two legs forward. Also, landing with one leg for-ward in the pumice overlay resulted in some directional instability and over-loading of the lower V-strut of the forward leg.4o


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z9 2 }

CONCLUDING REMARKS

Results of an experimental investigation have shown that the landing char-con-

, two long legs

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TABLE I.- SCALING LAWS

[A = Geometric model scale factor; P = Gravitational ratio, 6]Quantity Full-scalelunar model Scalefactor Earth-modelscale

LengthStress (yield strap) CTAcceleration a R PaArea A A2 A2AForce aA ? 1 2 ?2FMass F/a ? , 2 / p (A2/p)m

Velocity f a - 1 F P _ x f p _ A VTime V/a V /P U50 tInertia ml2 - A 4 / pAngular velocity 1/t C P 7 7Angular acceleration 1/t2 P / ? \

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Ve r t i c a l H o r izon t a l Fl igh t - p a t h P i t c h G e a r o r i e n t a t i o n ,Co n fi gur a t i o n Ru n speed, speed, angle, attitude, num b e r o f l e gs Sur fa c e c o n d i t i o n Commentsft/sec ft/sec de g de g fo r w a r dSy m m e t r i c a l 1 1 5 0 90 0 --- Flat, hard Ve r y l i t t l e r e b o un d .

fo ur p o i n t2 1 0 1 0 4 5 1 5 On e Flat, hard Undamped pitch oscillationb e t w e e n l e a d i n g a n d t r a i l i n g

p a d s a b o ut t w o s i d e p a d sd ur i ng s l i d e o ut .n s t e a d yp l a t fo r m .

3 1 0 10 4 5 0 T w o L e d g e d , h a r d Tw o t r a i l i n g p a d s i m p a c t i n i -(2-ft ledge) tia l l y o n l e d ge fo l l o w e d b yt w o l e a d i n g p a d s i m p a c t i n go n l o w e r l e ve l .t a b l esmooth slideout.

4 1 0 1 0 4 5 0 On e L e d ge d, h a r d I n i t i a l i m p a c t o n t r a i l i n g(2-ft ledge) leg.i d e p a d s m i s s e dl e d ge a n d e s s e n t i a l l y a l le n e r gy a b s o r b e d b y l e a d i n gleg.e a d i n g- l eg s h o c kstrut bottomed.

5 1 0 1 0 4 5 1 5 Tw o Fl a t, pum i c e o ve r l a y Stable.( 5 - i n . o ve r l a y)6 10 1 0 4 5 -15 Tw o Fl a t, pum i c e o ve r l a y Overturned.h o r t s l i d e( 5 - i n . o ve r l a y) th r ough pum i c e a n d s m a l ls h o c k - a b s o r b e r s t r o k e .7 1 0 1 5 33 0 On e Flat, pumice overlay Ve r y n e a r l y o ve r tur n e d a s( 21- i n . o ve r l a y) veh i c l e ve e r e d t o l e ft .8 10 1 5 33 -1 5 On e Fl a t, pum i c e o ve r l a y L o w e r V- s t r ut o f l e a d i n g l eg( 21- i n . o ve r l a y) buc k l e d a n d m o d ul e o ve r -

turned.m p a c t s i m i l a r t oh i t t i n g s o l i d o b s t a c l e .

Asymmetrical 9 1 0 10 4 5 -1 5 Tw o Fl a t, pum i c e o ve r l a y Stable.n c r e a s e d h o r izon t a lfo ur p o i n t ( 5 - i n . o ve r l a y) di s p l a c e m e n t t h r o ugh pum i c ea n d i n c r e a s e d s h o c k - a b s o r b e rstroke.

1 0 1 0 1 0 4 5 -1 5 Tw o L e d ge d, pum i c e o ve r l a y Stable.a t l o w e r l e ve l( 2- ft l e d ge, 21- i n .overlay)

Sym m et r i c a l 1 1 1 0 1 0 4 5 0 On e L e d ge d, h a r d M o d ul e r i gh t e d i t s e l f a ft e rfi ve p o i n t (2-ft ledge) slideout.table platform.

1 2 1 0 1 0 4 5 -1 5 Tw o L e d ge d, pum i c e o ve r l a y Stable.a t l o w e r l e ve l( 2- ft l e d ge, 5 -in.overlay)F "UI

TA B LE I I . - SU M M A R Y OF CH A R A CTER ISTICS OF TY P ICA L TEST R UNS


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Figure 1-63-553GEAR CONFIGURATIONS(a) S Y M M E T R I C A L 4 - P O I N T A R R A N G E M E N T .


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PLATFORM

~n m ^*~^p*o^" w ^"n^"^^~ ^TEST APPARATUS^

F L A T

LE08ED

Figure 3

HARD-SURFACE LANDINGSSYMM ETRICAL 4'P0|NTGEAR ONE LEG FORW ARDN TW O LEGS FORW ARDMAX. ACCEL-____o-------5`__^^^1`__^_^^^^-NNORMAL,2g units O2LONG.,gunits

-220ANGULAR,

RADIANS OSECL-20 OO55HORIZONTAL SPEED, FT/SECFigure 4


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l

" v*~* r ya po~n *^^ ^^ ^ `_USTSURFACE LANDINGSSYMMETR| CAL 4'PO|NTGEARSTABLE OVER- DEPTH,T U R NN.o^1^l} ~5NORMAL 2gunib O2LONG.,gunits

-220

ANGULAR.RADIANSSEC -20

OO5O5HORIZONTAL SPEED, FT/SECFigure 5DUST SURFACE LANDINGS

OVER TU R N CH A R A CTER I STI CS

20oSTAOLE vE-U R NPITCHSYM.4-P[oSYN. 5-PTAT[ 'ASYhl4'P[D E Go\^- 0o

o0050H OR I ZONTA L SPEED, FT/SECFigure 6


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AMMMMMMOW6. LUNAR TOUCHDOWN OPERATIONS AS AFFECTED

BY ABORT CONSIDERATIONSBy Gary P. Beasley

SUMMARY

A preliminary analytical study was conducted to determine the boundary orcurves for combinations of altitude and descent rate for which aborts

The results of the study established the minimum altitude threshold fornot possible.

INTRODUCTION

The final letdown phase of the lunar landing mission (the region from hover

define the boundary or "deadman" curves of rate ofthe combi-

determine the boundary curvesrange of lunar-excursion-module (LEM) characteristics andinto account.

SY M BOLS,F l ,F 2 ,F 3 ' F 4 r e a c t i o n - c o n t r o l e n g i n e t h r u s t a t v a r i o u s l o c a t i o n sAhrust of abort enginel t i tud eate of descento altitude at time o f a b o r to ra te o f descent at time of abortoment armResponse time

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Warth weight8itch angle0 0itch angle at time of abort0itching accelerationVEHICLE CHARACTERISTICS AND OPERATIONAL CONSIDERATIONS OF THE STUDYThe LEM control parameters taken into account in the study are shown in fig-ure 1 and consist of the thrust capability of the reaction-control system (RCS)engines (which could be throttled), the moment arm at which these engines fired,and the thrust capability of the main abort (ascent) engine. These quantitiesdetermine the reaction-control acceleration and the main deceleration capabilityavailable.

The pertinent conditions at the time of abort are the pitch angle, rate ofdescent, and altitude. Along with the control parameters and conditions, thestudy also included some operational considerations, one of these being that thepitch angle be zero or near zero before staging occurs and the abort (ascent)engine fires. This statement is not meant to imply that the abort engine must bevertical before staging and firing but was assumed for efficient use of abortengine thrust. This assumption also presents the more severe case to be studied.Other operational considerations were the decision time of the pilot and thestaging time of the LEM. It might be noted here that for the study an abort wasconsidered successful if the rate of descent of the LEM was halted before the LEMhit the moon.

BOUNDARY CURVES FOR AN ASSUMED LEM CONFIGURATION

In order to provide a first look at the type of results this study willproduce, some boundary curves were prepared for an assumed nominal LEM configura-tion. The nominal LEM parameters to be discussed and the boundary curve for thisconfiguration are shown in figure 2. As indicated in the figure, the maximumthrust level of the RCS engines was considered to be 75 pounds per engine at amoment arm sufficient to provide a pitching acceleration of 50er second squaredby using differential throttling of two engines. The thrust-to-earth-weight ratioof the abort engine was assumed to be 0.6 and the initial pitch angle of the LEMwas assumed to be 400t t h e t i m e o f a b o r t . T h i s a p p e a r s t o b e a r a t h e r l a r g epitch angle for the LEM to have at such low altitudes, but large angles such asthis might be encountered when establishing translation velocities as discussedin paper no. 4. ther considerations assumed for the nominal LEM configurationw e r e t h a t: ( 1 ) t h e L E M w a s r e q u i r e d t o b e p e r p e n d i c u l a r t o t h e l o c a l v e r t i c a lbefore the abort e ngines fired and (2) a 3- second combined pilot reaction and LEMs t agi ng t i m e w a s n e c e s s a r y fo r t h e a b o r t p r o c e dur e .

When the LEM configuration and initial conditions as well as the operationalconsiderations mentioned are taken into account, the study provides the boundarycurve fo r t h i s c a s e a s i n d i c a t e d i n figur e 2.

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^ x

The abscissa in figure 2 is the altitude of the LEM at the time of abort anddescent and

altitude for free fallvelocity of

moon's surface regardless ofcapability during

EFFECTS OF CHANGING PARAMETERS

To indicate the effect on the boundary curves of changes in the LEM param-ters and conditions, a preliminary study of some changes and their consequencess shown in figure 3. The first c urve on the right in figure 3 is the nom inalcurve mentioned in figure 2. The first two curves to the left of the nominalcurve show the effect on the nominal boundary curve of changing the abort engine

s e c o n d s . T h e s e c h a n g e s p r o v i d e l i t t l e i n c r e a s e i n a b o r t c a p a b i l i t y . T h eof having a pitch acceleration of 28.60/sec2

(0.5 radian per second squared) which was found to be desirable from a pilot'sstandpoint in the simulation study discussed in paper no. 4. This increase inast two curves are concerned with the problem of the pitch attitude of the LEM at

a n d fi r e d a t 40 0 to the vertical until the rate of descent is halted rather thanerecting the LEM and then firing. The second curve represents the case where theabort or survivable free-fall capability throughout the entire letdown phase. Theh e s a m e t i m e p i t c h i n g t h e L E M b a c k t o 0 = 0 p r o v i d e s a b o u n d a r y c u r v e c l o s e t ohe 0 0 = 0 curve.

Results from a piloted simulation study of lunar landings (paper no. 4) indi-c a t e d t h a t t h e s u c c e s s f u l c o m b i n a t i o n s o f h a n d h w h i c h o c c u r r e d t e n d e d t of a l l t o t h e r i g h t o f t h e 0 0 = 0 curve. In general, the combination of h and t e n d e d t o b e c l o s e r t o t h e 0 0 = 0 c u r v e a t a l t i t u d e s n e a r 4 0 0 f e e t t h a n a tthe lower altitudes. Thus, in case of abort situation, abort near an altitudeof 400 feet would have been more critical than an abort at a lower altitude.

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CONCLUDING REMARKS

Results of a preliminary study have indicated that it is impractical toreduce the lag time to below 1.5 seconds or increase the thrust level of theabort engine because of abort consideration alone since little abort capabilityis derived from large changes in these parameters. It can also be concluded thateven though relatively large changes in the no-abort region are obtained bychanging the pitch acceleration, it still does not provide the abort capabilityrequired. The results also show that in order to provide abort capability downto zero altitude, the vehicle must be made capable of firing the abort enginewhile tilted or, if this is not possible, it must be kept erect, a condition whichmight dictate the need to provide translation rockets on the LEM. Combinationsof these changes will also provide the abort capability necessary.

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ho I

LEM VARIABLES INVESTIGATED

Figure 1

NOMI NAL LE M CONFIGURATI ON100 - F = 7 5 L B P E R R C S E N G I N E

0 = 5 o P E R S E C 28 0 e 0 =4do

4 '

0

3 S E CR = 3 S E Ch0h o , - 00. 6W /

P S

T

N O - A B O R TO U N D A R Y4 0 R E G I O NF R E EA F E - A B O R T02 0 LF F A L LE G I O NA L L0

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2 0 0000 00 00 0 02 0 04 0 0h o , F T8 06 0h o , F P S402 00 5 S E CL L E M1 0 0E F F E C T S O F C H A N G IN G L E M V A R I A B L E S o --- --- 2Figure 352


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the indicated points. In performing most of the self-locomotive activities undernormal conditions, the body members move primarily in parallel planes; conse-quently, the test subject is able to perform the same activities in a more or lessn o r m a l m a n n e r e v e n t h o u g h t h e b o d y m e m b e r s a r e r e s t r i c t e d b y t h e c a b l e s . A s o m e -what oversimplified illustration of this point is that of a person walking down avery narrow hall.

Figure 2 illustrates the general layout of the equipment that was developed.The test subject is supported by a series of small cables attached to a light-weight crossbar which is in turn attached to a trolley which is free to move alonga monorail. A test operator or, possibly, a servo-controlled drive unit is usedto keep the trolley and, consequently, the cable directly over the test subjectas he moves either forward or backward. An inclined walkway runs parallel to themonorail track and represents the surface of the moon or the floor of the spacestation. The displacement of the walkway from directly beneath the track estab-lishes the inclination angle of the test subject, the supporting cable, and, con-sequently, the magnitude of the simulated reduced gravity. The case of zerogravity, or weightlessness, is produced with the walkway directly beneath thetrack or without the use of any walkway. It should be noted that this techniqueis useful for studies of the mechanics of self-locomotion or other dynamic prob-lems but is not suited to physiological or psychological studies of approximateweightlessness inasmuch as the internal organs are still subject to the earth'sgravity.

A photograph of the suspension system that was used in the development of amock-up of the simulator is shown in figure 3 he test subject's body is sup-ported at the head, the upper torso, the buttocks, and the calf of each leg byindividual suspension cables attached to the crossbar shown in figure 2. In somecases, depending on the type of task the subject wished to perform, the arms werealso supported just below the elbo w joint. Support of the lower leg proved to bethe most difficult problem but was achieved by use of a lightweight metal tubularframe that supported the leg from behind and permitted freedom of movement for theupper leg. The frame was supported by one of the suspension cables.

The mock-up of the simulator, part of which is shown in figure 3utilized amonorail system which had an elevation above the floor of about 40 feet and per-mitted a total cable length of slightly less than this distance. A small trolleywas used for this mock-up but for simplicity was not controlled by an operator ora drive system to keep it directly over the test subject at all times. Conse-quently, the test subject had to drag the trolley with him as he moved back andforth along the walkway. A 16-foot walkway was used for these preliminary testsbut was considered too short for practical testing purposes.

TESTS AND RESULTS

Several test subjects of various physiques and dressed in normal streetclothes were used to perform a series of locomotive tasks consisting of walking,running, jumping vertically, and climbing ladders, poles, and stairs both with

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of earth gravity and simulated lunar gravity were taken; also, the subjectson their ability to perform under both conditions.In general, the subjects had little difficulty in walking and running at

lunar gravity. The subjects were able to jump verticallystanding position in the earth gravity and could jump thefeet in the simulated lunar gravity; however, they experi-jump

The cable length of about 37 feet proved to be too short for these jumpingthe change in simulated gravityangle as the subject increased his distancetwice this length is recommendedwith the mock-upmass of the overhead trolley which interferedthis problem.

The subjects had no difficulty in climbing stairs and ladders at variouskeeping the legs properlythe lunar gravity

The test subjects could perform easily with no practice a number of gymnasticforward flips in the simulated lunarseconds to become accustomed to thethe body movements normally without loss of

CONCLUDING REMARKS

In conclusion, the developed technique for simulating reduced gravity appearsfor studying many aspects of the self-locomotive capabil-

form seriously limit the wearer's capabili-studies in which this technique is utilizedthe design and development of efficient space-station and

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useful auxiliary locomotive devices such as jump packs to assist the lunarexplorers. Finally, these test results should be helpful in planning the effi-cient logistic support for the lunar-exploration missions

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FORELF-LOCOMOTION STUDIESFigure 1-15 86REDUCED-GRAVITY SIMULATOR

MAN SELF-LOCOMOTION STUDIES

Figure 2-163+57


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TAKE-OFF,RENDEZVOUS,AND DOCKING


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8. SOME ASPECTS OF MAN'S VISUAL CAPABILITIES IN SPACEBy Jack E. Pennington

SUMMARY

Several studies have been conducted to determine a pilot's capabilities tois efficient in many areas, and in other areas visualbackup control techniques. This paper

INTRODUCTION

A number of studies of man's capabilities to control rendezvous have been5) show that manual control of rendezvous under

ities in a space environment.This paper summarizes the results of subsequent studies of man's visual

SYMBOLSeparation distance, ftlosure rate, ft/secisual angle, degininimum RENDEZVOUSOne technique for control of rendezvous is to b ring the angular rate of the

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then to control closure rate to effect a safe rendezvous. In s imulations uti-lizing this technique with visual r eferences for background and target, it wasfound that pilots could detect the target and establish the intercept course.However, at the time these studies were made, no data were availabl e to indicateto what accuracy the pilots could null the line-of-sight rate. Therefore, forthis information, a separate study was made to determine the smallest angularrate that a human pilot could be expected to detect. (See ref. 6.) The samebasic equipment used in the rendezvous simulations was used for this visual-acuity study.

An important parameter in the detection of angular rate is th e angularseparation between the target and the nearest reference star. The range ofseparation values between the target and the background used in this visual-acuity study was from zero separation (or superimposed condition) to an angle ofseparation of 60 milliradians. For this superimposed condition the pilot couldimmediately detect the rate of 0.1 milliradian per second. Results of thesevisual-acuity tests made with angular separations between the target and thenearest background star of 12.5 and 34.0 milliradians are shown in figure 1.This figure shows that at a separation distance of 12.5 milliradians, less than10 seconds is required for th e pilot to detect the desired rate of 0.1 milli-radian per second that was set forth in the pilot-control study. As shown inthe figure, the ability had deteriorated somewhat at a separation of 34.0 milli-radians. This reduced ability indicates that the target must be within 12 to15 milliradians of a reference star in ord er that the angular rate of 0.1 milli-radian per second be detected.

A study of star charts has shown that, on the average, a pilot can expect tohave a visible star (sixth magnitude or less) within 2 0 of the target. Thismeans that the pilot may find it necessary to delay the line-of-sight correctionuntil the target is in close proximity to a vis ible star or, alternatively, touse an optical aid either to superimpose the target on a star or to magnify andmake more stars visible. For instance, use of a 3 -inch telescope would permitan average density of 4 to 16 stars (11th magnitude or brighter) per squaredegree. This density would place the target within 12 milliradians of a starand permit detection of 0.1 milliradian per second. In reference 5 a study wasmade of a pilot's ability to determine the r ange and range rate from visualsightings. The analysis and simulation results indicate that it should be pos-sible to determine range and range rate from th e angular rate of the target.This technique depends upon development of an optical device which has not yetbeen demonstrated.Naturally, the first objective in rendezvous is to recognize and track thetarget. In future missions a beacon will be used for rendezvous practice andpossibly for target identification. In order to study the ability of an astro-naut to acquire and track a flashing beacon and to measure h is night adaptationlevel throughout the orbit, Langley Research Center (LRC) and Manned SpacecraftCenter cooperated in developing su ch a beacon and included it in the MA -9 flight.The beacon's trajectory prevented its being seen during part of the first orbitafter release, but on succeeding orbits Major L. Gordon Cooper periodically com-pared the beacon intensity with selected stars or with an onboard standard source.These data are still being evaluated.

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DOCKING

In the visual-docking maneuver the rate of change of size could be used to7. T h e s tu d y w as p r im ar il y concer ned

d r iv en in a cl os ed -l oop s ys tem for cl os u r e cu es . T h e

ystem. The results may be applied to the visual-contact portion of the docking.observer over the range of visual angles considered. The results, shown as aj e c t ) t h e b o u n d a r y o f t h i s S I S r a t i o . T h i s t h r e s h o l d w a s b a s e d o n a r e p l y t i m eof 2 s econd s d ictated b y tim e l ags inh er ent in th e tes t p r oced u r e. T h e figu r e isof interest because it shows that the maximum perception of closure occurs at vis-

0 to 90 0 . T h i s b o u n d a r ya g r e e s w i t h a n a n a l y t i c a l d e r i v a t i o n o f t h e r e l a t i o n b e t w e e n S a n d S . F o r t h i ss u b ject, it w as fou nd th at a r ep r es entativ e v al u e for th e cl os u r e th r es h ol d SMinSshould be able to judge the closure rate to about 0.15 feet per second from a dis-ance of 10 feet, a value which agrees with results of preliminary visual-docking

simulation studies conducted at LRC in which closed-circuit television was used.The results of these studies are discussed in paper no. 9 by Byron M. Jaquet andDonald R. Riley.

A test has been conducted inside the U.S. Navy 2,800-foot-long hydrodynamicaccurately estimate the separation distance to a target of known size with nocu es ex cep t th e ap p ar ent s ize of th e tar get. A fter a p er iod of d ar k ad ap tation,subjects were asked to estimate the range of several models of known size placedat r and om d is tances . T h e m od el s u s ed ar e s h ow n in figu r e 3. T h ey incl u d ed th r eed is ks , a tr iangl e, th r ee cyl ind er s s cal ed to 1 / 5, 1 1 0, and 1 20 th e s ize of th e

fl igh t. M od el s w er e p ainted b oth fl at w h ite and fl u or es cent or ange s o th at col oreffects could be investigated. Figure 4 shows the average distance judgments ofseveral observers for various configurations. The solid line in the figure rep-resents perfect estimates. Beyond 500 feet, average estimations were better thanexpected but with a tendency towards overestimating the range of the large objectsand u nd er es tim ating th e r ange of th e s m al l er tar get ob jects . A l l av er age es ti-mates (except for the balloon) were fairly accurate for distances from 500,to0 feet.

Although the average estimations for most models were accurate, the individ-ual estimates varied widely, and in many cases the average was accurate becauseindividual estimates "compensated." When the standard deviation of individualestimations was considered, it became apparent that estimates were reasonablyaccurate (20- to 25-percent err or) within ranges from 300 to 500 feet for areceding model, depending on its size, but estimates were accurate only within

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a compilation of recent research related to the apollo mission

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