STUDY OF AUTOMATED RENDEZVOUS AND DOCKING FOR ATS-V DESPIN

N7 1-29 6~ 1 AA~I~?J~B. R) ------U (PIES C I(NASA CRO RI RADNUBR (CATEGORY)

FINAL REPORT VOLUME 2 OF 2 I

SD 71-286, S Rproduediby-

INFORMATIONNA~OA TECHNIAL SERVICE U S Deportment of Commerce Springfield VA 22151 k SpaeDMaRon 1014NortAmerican RlodCl @l Space Division North American Rockwell

FINAL REPORT ON A STUDY OF AUTOMATED RENDEZVOUS AND DOCKING FOR ATS-V DESPIN

Volume 2 of 2 Volumes

Appendices

February 1971 Contract NASW-2136

Prepared for

NASA Goddard Space Flight Center

Greenbelt, Maryland ?:

SD 71-286 A1b Space Division @IN NorthAmencan Rockwell

PREFACE

NR has completed a study in accordance with the Statement of Work in NASA Contract NASW-2136, "Study of Automated Rendezvous and Docking for ATS-V Despin". Work was performed for NASA Headquarters under the technical direction of J. R. Burke and for the NASA Goddard Space Flight Center under the techni­ cal direction of J. Phenix. Their contribution and guidance to the study is gratefully acknowledged.

This final report, SD 71-286, presents the results of the study. It consists of 2 Volumes: Volume I - Text; Volume 2 - Appendices (Calculations and pertinent reports generated during the study). This is Volume 2 - Appendices.

SD 71-286 Space Division 9 NorthArnencan Rockwell

CONTENTS

Appendix A - Spin Axis Precession For Required Apogee Burn Attitude

Appendix B - Simulation Requirements Document

Appendix C - Piloted Docking Simulation Study

Appendix D - Post Contact Docking Analysis

Appendix E - Typical Procurement Specification

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V SD 71-286 i%Space Division North American Rockwell

APPEND IX A

SPIN AXIS PRECESSION FOR REQUIRED APOGEE BURN ATTITUDE

As described in Section 2 of the report, the RMU spin axis must be rotated through 1480 during the transfer orbit. The torque vector must be applied along an axis normal to the major axis of the transfer ellipse, normal to the spin vector in the plane of desired precession. Because of the spin, the torque is applied by means of periodic synchronized torsional reaction jet impulses. The angular velocity of the RMU about its spin axis, WN, will be between 40 RPM and 60 RPM. The sun sensor will be used to synchronize the cycling of the applied couples.

Since this precession will be initiated soon after perigee injection, an almost full propellant load will exist with a resulting nominal pressure of 170 psi. The corresponding mean-thrust level for intervals greater than 100 m-sec will be approximately 1.8 lb force per thruster. The distance between thruster pairs about the transverse axis is 54.4 inches.

Thus, the expected nominal torque Tnom during the time the couple is applied is:

T = 4x 1.8 = 8.1ft-Ibs nom 12

Consider the precession scheme where four sets of two actuators are sequenced to produce couples, -r/2 radians of spin apart. This sequence could cycle let numbers (1-8), (7-6), (5-12), (2-10) in that order (seereaction jet logic, Section 7). Note that any wobble-countering couple required during the time the precession takes place can be obtained from the respective alternates of the above sequence, namely (3-9), (3-11), (4-11), (4-9), without interferring with the precession cycle. Let 9 th denote the half-angle of spin over which the couple acts.

Let & (A0) denote the precession angle obtained per revolution of the RMU. Then, gth

8 Tno m Wx Cos (wxt)dt= 8 Tnom Sin L cXxlx Ix x

Here

Tnom = nominal torque 8.1 ft-lbs

A-1S SID 71-286 Space Division ?D North Amencan Rockwell

2 of inertia about spin axis - 57 flug-ft Ix = moment

w = spin rate = 60 rpm nominal and 40 rpm minimum

9 So that from (1), 6 (.) = .0288 (Sin th ) rad/rev x 60 rpm

9 .0648 (Sin th) rad/rev (OWx= 40 rpm)

The number of .revolutions required to achieve a given precession angle AO is then:

.A0A(Nrev)A$(the total number of precession increments =4 Nrev)

Thn total time required to complete a precession thru AO is:

tpr =(N)ev 2 Tr 7 (. x 4 InomSin 0th

So that for (A$) = 148 deg = 2.58 rad

(Nrev) = 89.6 (for o =60 rpm) re14 8 0 = 8- x,

= 39.8 (for wx =40 rpm) Sin 9th

Similarly:

(tprpr)1480 Sin89.6 gthx' sec (for co =60 rpm)

- 59.7__ sec (for w =40 rpm) Sin 9th x

The propellant weight required to achieve a given (A $) is:

2 W = F tpr 8 gth . , where Is= spec impulse of propellant prpsp = 200 sec (assumed) Fa = force per actuator = 1.8 lbs (assumed)

A-2 SD 71-286 Space Division NorthAmerican Rockwell

So that for A 1480

(Wprop) 1480 = 2.06 )bs1h for (& = 60 rpm)

= 1.37 (St 9 th Ibs for = 40 rpm

A plot of (tpr)148a and (Wprap) vs 9 th is shown in Figure 2-10. A near optimal

choice is 9th = 450* Here (tpr) is a near minimum with no appreciable cost in propellant weight increase.

For the purpose of this analysis, the precession trhusting half-angle is assumed fixed at gth = 450 for the open loop precession to achieve A,0 = 148 degrees.

The following parameters then result for the 1480 open loop precession

W =60 rpm W =40 rpm x x Time to complete A '=1480, (tp) 148o 126 sec 84 sec

Total number of precession increments,

N I =4N rev- 507 225

Magnitude of precession increments =

1/4 8 (A ) 5.08 mr 11.45 mr

Impulse bit time interval per increment 250 m sec 375 m sec

Precession Error Analysis

The error in orientation after precession consists of a component in the precession­ plane, Ee and a pointing error normal to the precession plane, e . It will be con­ venient t6 represent tkese two error angles as vectors, a permissible approximation since these will be small angles (e8 < e4, < 1/4 rad). All error quantities are assumed 3 a values.

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The individual contributors to this attitude error are identified as follows: Pal 1. Error due to POLANG Attitude Determination - E0 The value of this

total error after a minimum of 1.5 hours of ground station POLANG smoothing can be assumed at 0.50 or = 8.7 mr. Since the POLANG sensitive axis is nearly normal to the spin when the measurement is required, a propagation coefficient of 1 .0 is used so that

Pal E =8.7 mr

No capability for measurement in the 9 direction is provided. SSA SSA 2. Error due to sun sensor attitude determination, E E ) Since the

sun direction = 450 with respect to the nodal line and the apogee spin orientation (where measurement is required) is at -22 degree from the equatorial plane, the sun sensor error propagation coefficient is approximately 1/Sin 500 = 1.3 for both 9 and $ directions. With an expected instrument error of 0.5 degree, this results in: SSA SSA = = 11.3 mr

3. Error due to sun sensor error in synchronizing precession torque axis, SSP E If the precession vector is constantly offset by 0.5 due to sun

sensor error (pessimistic), the accumulated error normal to the precession plane is (0.5/57.3) A$,

So: SSP SSP =0, 6 = 0.00873 (A (P ) mr

4. Errors due to the actuator pairs supplying the precession torque increments. These error contributions are best described as four individual contributions:

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(a) Impulse bit repetition error due to the variation in the average force from given actuator from the expected mean for any given increment. The expected maximum repetition error is 6%. Since two independent actuators form a given couple, the resulting torquing uncertainy is: 1 x 06 = .0415,;

and is averaged over the total number of increments, N1 for a given A,

Further, this error contribution is confined to the O-plane only. Thus, the impulse bit repetition error becomes

afr afr 60 S -0,_ A mr

(b) Impulse bit bias error for a given actuator is assumed as 4% max. This bias is correlated for a given actuator with the eight individual jets independent of each other. Again, this contribution is confined to the $-plane so that:

afb afb 40 E9 01 E = (A0) mr.

(c) Impulse bit centroid repetition error. This contributor is confined to the 9-plane since centroid shift timing for a given torque increment is equiva­ lent to a shift in the precession axis. The magnitude of the max centroid shift from pulse-to-pulse is assumed at 6%. It results in precession axis tilt of [6 (0th)] where 9 th is spin angle corresponding to the pulse width - 450 . This error is averaged over the total number of pre­ cession increments. The term arises from the two jets used for a given increment. Thus, the impulse bit centroid repetition error becomes: acr acr 47.2 ( A mr, E0 -0 2WI

(d) Impulse bit centroid bias error. This contributor again is confined to the 9 plane. It is correlated for a given actuator and averaged over the eight ihdependent jets used per spin cycle. The maximum magnitude of the centroid bias is 4%. Thus

acb acb 31.4 ()mr, 0

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(5) Torquing Increment Quantization error. The maximum value of this con­ tributor is simply 1/2 (precession increment), so that

aq aq E =0, E 2.5 mr (for wx=60rpm)

5.7 mr (for w = 40rpm) x

(6) Torque Integration Error. During a given precession increment, the torque vector rotates through 900 about the spin vector. The time integral of these torques is in the central (average) direction if the angular shift of the spin axis is negligible. If the spin axis shifts substantially during the increment, the effective precession axis shifts from the theoretical pre­ cession axis (due to the fact that Euler angles are integrals of orthogonal axis rates rultiplied by transcendental functions of the Euler angles them­ selves). However, since for the current systems the spin axis changes by less than 10 per each increment, the shift of the effective precession axis is considered negligibly small by comparison with other sources considered having the same effect (#3 and #4 c, d).

The expected maximum error in attitude effort precessing the RMU open loop through 1480 in the plane normal to the nodal line can now be computed for the worst case condition of Cx = 40 rpm. The base parameters are:

AO= 1480 = 2.58 rad

N I = 225, (number of precession increments)

A data smoothing time of, say two hours, is required to reduce the POLANG errors to the value given; then making use of error sources I through 5, there results (errors in mr):

Pol E =8.7

SSA SSA E = 11.3 e0 =11.3

Total attitude alt att Assumes statistical determination E = 11.3 E = 6.9 combination of error contri- redundant measure­ bution ments

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afr afr C =0 C =7.3

afb afb = 0 C =36.5 9

acr acr E= 5.8 E =0 o 0

acb acb E = 28.6 E = 0 9 0

aq aq e = 0 E =5.7 o 0

SSP SSA = 22.4 E = 0

Total precession caused error a a contribution C = 36.8 E = 37.7

Thus, the combined maximum pointing error including the attitude error after two hours of smoothing time following completion of the open loop 1480 precession is given as: art 2 a2 att2 a2 (A0)2 = kE ) +(E ) + (E ,) + (E )]1/2=54.3mr=3.13deg.

A secondary ,precession correction of this resulting pointing error can then be accomplished with base parameters for the worst case condition as:

= 3 13 = (A 2 =) deg. .0543 rad

(N1)2 = 5 (number of precession increments)

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The resulting pointing error after this second open loop precession through 3.13 degrees becomes (errors in mr): Pol E =8.7 0

SSA SSA E= 11.3 E = 11.3 0 $

Attitude aft att Determination e =11.3 E 6.9 Error 9 , Contribution

afr afr 0 e =1.1

afb afb =0 g0 E =0.7

acr acr =0.8 =0

acb acb =0.6 E =0

aq aq 6=0 E =5.7

SSP SSA = 0.07 E = 0 o 0

Total Precession a a Caused Error 1.0 E = 5.8 Contribution 9" $

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The combined maximum pointing error after completion of the open loop correction for A0 2 =3.13isthen

I (A0 3 =14.5 mr = 0.835deg

As shown above, this value is due almost entirely to the attitude determination uncertainty so that no significant further improvement can be expected from another precession correction of this measured value.

Accordingly, attitude control sequence to achieve the apogee firing orientation is as shown in the time-line of Figure 2-11. The indicated second correction at five hours can include increments reflecting measured orbital errors (tracking inputs) to reduce to some extent subsequent AV requirements.

A-9 SD 71-286 9111 Space Division North American Rockwell

APPENDIX B

SIMULATION REQUIREMENTS DOCUMENT

SD 71-286 #Ikk SpaceDivision 9I NorthAmencan Rockwell

FOREWARD

This simulation requirements document vas prepared in response to Remote Maneuvering Unit (RMU) contract NASW-2136 by G. 0. Mount, and D. B. Wurzburg. The objective of the document is to describe the RMU pilot-in-the-loop docking simulation, its equations, special effects, and data requirements to aid in the preparation of the simulation facility. The efforts of Messrs. W. Lowe, B. Morris, 0. Hayakawa, and N. Svensrud in developing simulation hardware requirements are expressly appreciated.

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SD 70-5WO 'lk Space Division North Amencan Rockwell

CONTENTS

SECTION Page

1. 0 MAN-IN-THE-LOOP DOCKING SIMULATION ...... 1 1.1 Introduction ...... ± 1. 2 Objectives...... 4 1.3 Scope...... 4 1. 4 Simulator Configuration...... 7

2.0 DEFINITIONS ...... 13 2. 1 Definition of Input Data ...... 13 2. 2 Simulator Nomenclature ...... 13 2. 3 Definition of Abbreviations ...... 13

3.0 SIMULATION EQUATIONS ...... 27 3.1 General Description ...... 27 3.2 Initial and Final Conditions ...... 27 3.3 Axis System and Rotation Order ...... 29 3.4 Mathematical Model ...... 29

4.0 SPECIAL EFFECTS EQUIPMENT ...... 37 4.1 Model Requirements ...... 37 4.2 Pilot Command Console Requirements...... 2 4. 3 Visual System Requirements...... )

5.0 PROBLEM CONTROL ...... 7 5.1 Responsibilities ...... 17 5. Z Simulation Accuracy Requirements .)...... 19 5.3 Simulation Checkout Requirements ...... 52 5.4 Preliminary Piloting Procedures ...... 6o 5.5 Test Plan and Schedules ...... 62

6.0 DATA ACQUISITION AND REDUCTION...... 69 6. L Output Data Requirements ...... 69 6. 2 Data Reduction Equations ...... 69 6.3 Data Records . .o...... U9 6.4 Quick Look Documentation ...... 70

-P5E NOTPI

SD 70-5cc 9 Space Division NorthAmerican Rockwell

ILLUSTRATIONS

FIGURE PAGE

1.1.1 RMU/ATS-V 2 1.1.2 ATS-V Orbital Status 3 1.3.1 Pilots View. 6 1.4.1 RMU/ATS Precontact Docking Simulation 8 1.4.2 GPVS Transport and Gimbal System 9 1.4.3 View of the ATS at Pilot's Monitor . 10 1.4.4 RMU Simulation System Block Diagram. 11 2.1.2 Rotational Controller (Yaw Axis) 20 2.1.3 Rotational Controller (Pitch Axis) 21 2.1.4 Rotational Controller (Roll Axis) . 22 2.1.5 Rotational Controller Rotation Rate Versus Stick Displacement. . . 23 3.1.1 RMU/ATS-V Simulation Axis System 28 3.3.1 Active Vehicle Rotation Order 30 3.4.1 Math Model . 35 4.1.1 ATS Model . 38 4.1.3 Spinning Docking System Model (Side View) 4o 4.1.4 Spinning Docking System Model (Front View) 41 4.2.1 Pilot Control Console . .13 4.3.1 Visual Display System - 45 4.3.2 Derotation Optics System 46 5.5.1 RMU/ATS Docking Simulation Schedule 65 6.4.1 Quick-Look Data Sheet 71

vii SD 70-5b0 91b Space Divsion North Amencan Rockwell

TABLES

TABLE PAGE

2.1.1 Definition of Input Data 14 2.2.1 Simulation Nomenclature . 17 2.2.2 Definitions of Statistical Data Reduction Symbols. . 24 2.3.1 Definition of Abbreviations 25 3.2.1 Initial Conditions 27 5.2.1 Rig Performance . 49

5.2.2 Gimbal Performance . . . .50 5.3-1.3a RMU Checkout Maneuvers (Position Check) 53 5.3.1.3b RMU Checkout Maneuvers (Velocity Check) 54 5.3.1.5a RMU Jet Checkout (Translation) 55 5.3.1.5b RMU Jet Select Checkout . 56 5.5.1 RMU/ATS Docking Test Plan 65 6.1.1 Recorded Data. . . 74 6.2.1 RMU Docking Simulation Results 72

x SD 70-580 ' Space Division North American Rockwell

SUMMARY

A man-in-the-loop, six degrees of freedom, docking simulation study is to be conducted to investigate the pilot/control system/visual system interface effects on the precision of angular and translational positioning of the RMU with respect to the ATS-V spacecraft during the final phase of docking. This document contains a complete description of the simulator requirements, problem control and data acquisition requirements.

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1.0 MAN-IN-THE-LOOP DOCKING SIMULATION

1.1 INTRODUCTION

The purpose of this document is to describe all the requirements of a man-in-the loop docking simulation necessary to establish pilot generated docking contact conditions (velocities, attitudes and miss distance) between the Remote Maneuvering Unit (RMU) and Applications Technology Satellite-V (ATS-V) as shown conceptually in Figure 1.1.1.

The ATS-V vehicle is presently in a stationary orbit- 20,000 nautical miles above the earth oriented as shown in Figure 1.1.2. The satellite is spinning in a stable condition at 73 RPM about its longitudinal axis. The spin is undesirable because it prevents deployment of ATS experiments. NR has proposed (Reference 1) using an unmanned vehicle with a spinning dock­ ing system to latch on, rigidize, and despin the ATS-V. The unmanned vehicle is to be pilot controlled from the ground through a television visual system.

During the docking maneuvers, the pilot must be able to see his target, and control the vehicle he is flying with extreme precision. Precision in translation and rotational alignment is required throughout the maneuver, but is especially critical at the time of contact between the two vehicles.

The ATS-V docking maneuver presents an unusual pilot task in that it is complicated by the following factors:

1. The ATS-V is spinning and requires a spinning docking system to be synchronized to null relative motion at the docking interface. The spinning docking system generates a precession 90 derees to the intended angular change of the vehicle when an external con­ trol torque is applied.

2. The ATS-V has its longitudinal axis pointing 90 degrees to the sun's rays such that the docking structure is unlighted.

3. Command signals from the ground and visual response from the vehicle to the ground take time to cover the 20,000 nautical mile distance. This timeJl3ag will make the control system response feel sluggish to commands. If the lag is large enough, the pilot commands could be 180 degrees out of phase with vehicle motion and tend toward instability. Line-of-sight communicat'ion via the RMU high gain antenna limits attitude excursions in pitch and roll to less than + 6.idegrees.

-1- SD 70-580 'ATS-V DOCKING RING

STATIONARY RMU

'J

' \

o cn SPINNIGATS-V 2 z"001 SPINNING DOCKING SYSTEM

Figure 1.1.1 RMU/ATS-V Docking Maneuver Direction of Relativ Motion for Docking- ATS-V ORBIT PATH

SPIN AXIS NORMAL TO ORBIT PLANE

* ORBIT: SYNCHRONOUS INCLINATION: +2.20 LONGITUDE: 2450 n IS 0 zw 0 0 Figure 1.1.2 ATS-V Orbital Status . C

2w Adlb Space Division ETV North Amencan Rockwell

4. The ATS-V spin axis is 1.6 degrees off of its geometrical center line such that, when perfectly aligned with the spin axis, the pilot will see a wobble of the spacecraft.

1.2 OBJECTIVES

The objective of the pilot-in-the-loop simulation study is to provide statistical data describing the docking contact conditions; that is, re­ lative velocities, attitude rates, attitudes and miss distances, between the RMU and the ATS-V. These data serve three purposes: they will permit evaluation of docking limitations and capabilities, tuney will provide the docking and visual'systems with design criteria, and they will aid in defining spacecraft control system requirements.

1.3 SCOPE

The scope of the RMU/ATS-V pilot-in-the-loop docking simulation -includes evaluation of piloting procedures, control systems, pilot console displays, instruments, and RMU visual systems. A rationale of each evalua­ tion and the method of analysis is as follows:

Piloting Procedures - A baseline piloting procedure will be recommended to each pilot for evaluation. Modifications to the baseline will be accepted if compatible with all other mission requirements and agree­ able to other pilots. Each pilot will undergo a learning process I:-ic!. will tend to establish his "standard" of performance of' the piloting procedure. An indicator of when the learning curve reaches a standard is when fuel consumption remains relatively constant for a number of maneuvers. The purpose of the "standard" is not to generate competition among the pilots, but to stabilize the statistical data of docking con­ tact conditions and reduce data scatter. The recommended baseline piloting procedure is outlined in paragraph 5.4.

Control Systems - A baseline RMU control system will be defined based on the control power requirements established during Apollo docking by scaling down jet thrust to the RMU mass properties. Variations of rotational control system sensitivity (i.e. rate commands vs. control­ ler deflection) and translation control power (i.e. translation com­ mand acceleration) will be evaluated. The basuline control system identified in paragraph 5.4 will be used throug:.oat the piloting pro­ cedures learning process primarily to save time lut also to establish a control system that should be close to optimun early in the study. The baseline control system, however, may be motified by the results of variation of control systems parameters. This portion of the simu­ lation study will be accomplished immediately followin, tLe "learning" process.

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Pilot Console Instrumentation and Displays - The pilot console will incorporate a number of instruments and displays which Will provide as much information on the dynamic state of the RMU, separately and with respect to the ATS-V, as required for docking. Much of the instrumentation will be redundant. Evaluation of these displays by the pilots will occur continuously throughout the simulation study without devoting a special set of maneuvers to the task. Suggestions for improvements will become design requirements if agreed upon by other pilots and compatible with other mission constraints. If practical, the improvements will be incorporated in the 'imulation. The baseline instrumentation and displays are defined in'paragraph 4.2.

Visual Presentation - The RMU will incorporate a closed circuit tele­ vision visual system which will permit the pilot to see the motion of the ATS-V with respect to the RMU as the RMU approaches for docking. ,The visual system will include'a method for visually stopping the rota­ tion of the ATS-V. The P14U docking system will also be-visible, as shown in Figure 1.3.1 so that it can be indexed to the ATS-V docking ring while both are actually spinning although visually stopped. Both the actual image and the visually stopped image contain informa­ tion that can be used for docking. Evaluation of this system, in addition to lighting intensities, and possibly a stero-optical system, will be conducted depending on the results of lab tests. A baseline visual system is defined for all previous simulation runs in paragraph 4.3.

Baseline Configuration Runs - When evaluations of procedures, control systems, instrumentation, and visual presentation have.been completed and recommended system changes have been incorporated where possible, a final set of runs will be conducted to gather statistical data on the baseline configuration.

Data Analysis - The simulation runs will be evaluated from three basic data sources; pilot opinion (de-briefing), fuel consumption, and dock­ ing conditions. The data from each run will be sammarized)on a "quick­ look" data sheet. At the end of a particular section of simulation runs both pilots and engineers will summarize the block of runs in a data summary report. A sample "quick-look" data sheet is shown in Section 6.0.

The Data Summary will contain histograms of docking contact conditions, and fuel consumption as well as a summary of'pilot opinion as to the effects of the configuration flown.

A final report consisting of system requirements defined by pilot opinion and numerical statistical data resultin, r'rom the simulation will form the nucleus of the RMU dodking system design requirements.

The statistical ahalysis' equations to be used in simulation data reduction are defihed in Section 6.2.

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a'

t) a A DOCKING LATCHESo ,PROBES 11 tGUIDES ON DOCKING 0

Figure 1.3 a Pilots View Space Msion ik NorthAm Rockwell

1.4 SIMULATOR CONFIGURATION

The visual simulator required to accomplish pilot-in-the-loop studies of the precontact docking maneuver of the RMU/ATS-V is presented schematically in Figure 1.4.1.

A television monitor command center, configured as a pilot console, accepts pilot commands through a set of translation and rotation hand controllers. The commands are transmitted to an analog computer.

The RMU and its control system respond to the pilot commands through the equations of motion. Relative motion of the RMU with respect to the"ATS-V are computed using transformation matrix equations. The resulting relative motion signals are sent to the GPVS Transport (on which a T.V. camera and an active model of the RMU spin cage are mounted) and to a gimbal system supporting the spinning ATS model in proper order to provide six degrees of freedom relative motion. The GPVS Transport and gimbal system are shown in Figure 1.4.2.

The television camera moves as if it were the RMU. The television image of the spinning ATS model and visible portions of the spin controlled RMU docking system are transmitted to the pilot console TV monitor.

The pilot closes the loop by responding to the visual cues of posi­ tion and motion by commanding attitude and translation changes as required to effect alignment and closure to docking contact. Special effects such as side lighting of the sun, signal/distance time lags, visible portions of the RMU docking system, and gyro cross coupling motion of the RMU caused by the spinning docking system all lend complete realism to the pilot/control/visual image interface. Figure 1.4.3 illustrates the view at the pilot monitor.

Realism is enhanced by the fact that this is a "fixed base" simulation of a fixed base control problem; that is, the pilot and pilot console are not required to move in a remote control situation (as compared with airplane simulators which are sometimes required to move and pro­ vide the feeling of motion to a pilot).

The above is a general description of the JMU/ATS docking simulator. Detail requirements of each component shown in Figure 1-5 are established in the following sections: Simulation Equations Section 3. 11 ATS and Docking System Model Section 1i.1 Pilot Console and Instruments Section 4.2 TV System and Lighting Section 4.3 Camera and Rig Section 5.2 Simulator Checkout Section 5-3 Data Requirements Section 6.1

775 SD 70- 5co T iwrRV(SUAL- PISPLAYf'lL qsr

Sy-em WPpfGI(NLSSrfl rUorR

TV0ME­

0

rcur,vuA z ~ PZCPJAT D)KNC,5MLTC! 0

Figure 1.4.2 GPVS Transport and Gimbal System

-9- SD 70-580 Figure 1.4.3 View of the ATS At The Pilot Monitor -10- SD 70-580 LOTGIC Discretes DT SLatch Index GCjDT ANALOG COMPUTATION ACQUISITION PpRange +A. & RanAttitude e Rate Simulated Control System -p PIO OSL Angular PILOT CONSOLE . Vehicle Dynamics Alignment TV Monitor TIME Spin Cage Controls DELAY Thrust shapping Angular Optics Controls T/C ELECTRONTCS Rate Translational Control . Fuel Accounting Rotational Control R/C 220+5.0 . Linear SRelative Attitude Milliseconds • Coordinate Transforma- Velocity Rangetions * Range Rate S'c tisn Index pLatchPinning . GPVS Drive Equations Distance I0/C mage . Pilot Display Drive . Fuel De rotation a) Optics De-Spun H Electronics Image % ELECTRONICS DEROTATION OPTICS * TV Monitor VISUAL DISPLAY SYSTEM * Dove Prism Spinning Image * 500 Line Camera 3 DOF Transport * Dove Prism Servo Drive RMU Docking Cage * TV Camera * ATS Model System T/C = Translation Commands * 3 DOF Gimbal R/C = Rotation Commands ZCD S/C = Spin Commands g O/C = Optics Commands B

FIGURE 1.4.4 RMU SIMULATION SYSTEM 0

BLOCK DIAGRAM -11­ Space Division 'D North American Rockwell

2.0 DEFINITIONS

2.0 DEFINITIONS

The purpose of this section is to specify all input numbers to be used in the computer as well as define all symbols, abbreviations and nomenclature used in the RMU/ATS docking simulation.

2.1 DEFINITION OF INPUT DATA

Input data symbols, value, definition and units are presented in Table 2.1.1. Figures 2.1.2 through 2.1.5 illustrate rotation hand con­ troller characteristics.

2.2 SIMULATOR NOMENCLATURE

All symbols used in the math model which are dependent variables are defined in Table 2.2.1. Table 2.2.2 defines symbols used in the final data reduction equations.

2.3 DEFINITION OF ABBREVIATIONS

Table 2.3.1 defines some of the abbreviations and acronyms used to describe various functions found in the simulation.

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TABLE 2.1.1 DEFINITION OF INPUT DATA

SYMBOL VALUE UNTITS DEFINITION

M 15.5 Slug RMU Mass

X 47-9h Slug-Ft2 EMU Roll Moment of Inertia

Iyy 44.2 Slug-Ft2 RMU Pitch Moment of Inertia

izz 46.1 Slug-Ft2 RMU Yaw Moment of Inertia 2 IXy 0.0 Slug-Ft

Iyz 0.0 Slug-Ft2 RMU Cross Products of Inertia

2 IXy 0.0 Slug-Ft J 2 0.5 Slug-Ft fDocking Cage Roll Moment of Inertia

AX 0.0 Feet Distance From RMU X-Jet Centers to X-C.G.

Ay 0.0 Feet Distance From RMU'Centerline to Y-C.G.

AZ 0.0 Feet Distance From PMU Centerline to Z-C.G. A R45 X -2.5 Feet Distance From l14U X-Jet Center to Camera Focal Pt. A R45 Y 0.0 Feet Distance From RMU Centerline Y-Camera Focal Pt. A R45Z 0.0 Feet Distance from RMU Centerline Z-Camera Focal Pt. A R43X 5.8 Feet Distance From RMU X-Jet Center to Docking Plane A R03Y 0.0 Feet Distance From RMU Centerline to Docking Plane A R43Z 0.0 Feet Distance From EMU Centerline to Docking Plane

H 2.21 Feet Jet Moment Ahn Along RMU X Axis

G- 1.54 Feet Jet Moment Arm AlongR U Y and Z Axes

OT 73.0 RPM ATS Roll Rate

-14-

SD 70-580 * Space Division North Amencan Rockwell

TABLE 2.1.1 DEFINITION OF IPUT DATA (continued)

SYMBOL VALUE UITS DEFINITION

T 1.5 Lbs Single RMU Jet Thrust

Isp 210.0 Sec Specific Impulse of RMU Jets 7OFF .038 See Jet Thrust Tail Off Time

TON .038 See Jet Thrust Rise Time t mi .027 See Jet Minimum Burn Time

.010 See Jet Solenoid Delay Time

To Unity None Jet On Command

d90 0.63 None Switch Gate Deadband, All Axes

OL +45.0 Deg Gimbal Angle Saturation Limit, All Axes

h +4o.o Deg Gyro Desaturate Switch Hysteresis

HG 1.2 Lb-Ft-Sec 2 Gyro Momentum, fll Axes

B .24 Lb-Ft-Sec DPG Feedback Damping Gain, All Axes 2 IG .00024 Slug-Ft 2 Gyro Transverse Moment of Inertia, All Axes

TG .0209 Lb-Ft Maximum DPG Torque Command, All Axes

6r 1.0 Deg/Sec Maximum Angular Rate Command, All Axes

a 0.22 Sec Time Lag for Ground Signal to Reach RMU and Back

6 TB 5,o Deg Translation Controller Breakout y z Deflection

FTB y,. 2.0 Lbs Translation Controller Breakout Force

-15-

SD 70-580 Space Division 90 NorthAmencan Rockwell

TABLE 2.1.1 DEFINITION OF INPUT DATA (continued)

SYMPOL VALUE UNITS DEFINITION 0.4 Inches Translation Controller Breakout 8 TBX Deflection

Kv 0.0167 RPM/DEG Docking Cage Spin-Up Vernier Sensi­ tivity

TSC 2.0 Lb-Ft Max Spin Cage Torque Command

.231 Rad/Sec Spin Command Break Point

40 73.0 RPM Spin Cage Switch Command

K0p 0.0083 RPM/DEG Despin Optics Vernier Sensitivity 4OP 36.5 RPM Despin Optics Switch Command

RTIPx +1.33 Ft Distance From ATS C.G. to focking Interface

RTI2Y 0.0 Ft Distance From ATS C.G. to Docking Interface

RT12z 0.0 Ft Distance From ATS C. G. to Docking Interface

For Rotational Controller Characteristics, See Figures 2.1.2 through 2.1.5.

±l6- SD 70-580 Space Division North Amencan Rockwell

TABLE 2.2.1 SIMULATION NOMENCLATURE

SYMBOL RANGE UNITS DESCRIPTION T R 23X 0-160 Feet X - Dist. Between Docking Planes, ATS Axis T R 23Y +20 Feet Y - Dist. Between Docking Planes, ATS Axis

R 23Z +20 Feet Z - Dist. Between Docking Planes, ATS Axis

RTI 3X 0-160 Feet X - Dist. Between ATS C.G. & RMU Docking Plane. ATS Axis T R 13y +20 Feet Y - Dist. Between ATS C.G. & RMU Docking Plane, ATS Axis T R 13Z +20 Feet Z - Dist. Between ATS C.G. & RMU Docking Plane. ATS Axis

RT15X O-160 Feet X - Dist. Between ATS C.G. & RMU Camera Focal Point in ATS Axis T RT 5Y +20 Feet Y - Dist. Between ATS C.G. & RMU Camera Focal Point in ATS Axis

RT15Z +20 Feet Z - Dist. Between ATS C.G. & RMU Camera Focal Point in ATS Axis

RT14Xo -160 Feet I.C.X - Dist. Between ATS C.G. & RMU C.G., ATS Axis

RTl4Yo +10 Feet I.C.Y - Dist. Between ATS C.G. & HWU C.G., ATS Axis

RT 4zo +10 Feet I.C.Z - Dist Between ATS C.G. & RMU C.G., ATS Axis v 14x -1.0 FPS X - Velocity Between C.G.'s, ATS Axis

VThY +-5 FPS V - Velocity Between C.G.'s, ATS Axis vTIhz +-5 FPS Z - Velocity Between C.G.'s, ATS Axis

Fx +3.0 Lbs Total Force, X-Direction RMU Axis Fy +3.0 Lbs Total Force, Y-Direction RMU Axis

-17- SD 70-5W Sce Division NorthAmencan Rockwell

TABLE 2.2.1 SIMULATION NOMENCLATURE (Continued)

SYMBOL RANGE UNITS DESCRIPTION

Fz +3.0 Lbs Total Force, Z-Direction MU Axis

WTOT Unk Lbs Total H'1 Propellant Used

WTR Unk Lbs Translational Propellant Used

WR Unk Lbs Rotational Propellant Used

A - Jet Pod Identification

B - (See Location with Respect to

C - - BMU on Fig. 3.4.1 of Sec. 3.0)

9C Unity - On Command, Pitch Jets

0C Unity - On Command, Roll Jets kC Unity - On Command, Yaw Jets

Xc Unity - On Command, X-Jets YC Unity - On Command, Y-Jets

ZC Unity - On Command, Z-Jets e" Unity - Roll Jet Signal, Solenoid " Unity - Pitch Jet Signal, Solenoid Q 41 Unity - Yaw JetSignal, Solenoid

'60 +6.0 Deg/sec Pitch Gyro Gimbal Rate "60 +6.0 Deg/sec Roll Gyro Gimbal Rate

'61, +6.o Deg/See Yaw Gyro Gimbal Rate

60 +50.0 Degrees Pitch Gimbal Angle

6 +50.0 Degrees Roll Gimbal Angle 6'!' +50.0 Degrees Yaw Gimbal Angle .a 2 V14 +.25 FPS Inertial Accel. X-Direction, IMU Axis

-18-

SD 70-580 Space Division North Amencan Rockwell

TABLE 2.2.1 SIMULATION NOMENCLATURE (Continued)

SYMBOL RANGE UNITS DESCRIPTION

.a+25 Inertial Aee. Y-Diction, BMU Axis

V14Z +25 Fl'S2 Inertial Accel. Z-Direction, RMU Axis

9 +10.0 Degrees RMU Pitch Inertial Attitude

+10.0 Degrees RMU Roll Inertial Attitude

+10.0 Degrees RMU Yaw Inertial Attitude dx +2.0 Deg/Sec RMU Roll Rate. Body Axis

y +_2.0 Deg/Sec BMU Pitch Rate, Body Axis

'kz +2.0 Deg/Sec RMU Yaw Rate, Body Axis

6 X +7.0 Deg/Sec 2 BMW Roll Accel Body Axis

6y +7.0 Deg/Sec2 FMU Pitch Accel. Body Axis

6Z +9.0 Deg/Sec2 EMU Yaw Accel, Body Axis MX +5.0 Ft-Lbs Total Moment About BMU X Body Axis

My +5.0 Ft-Lbs Total Moment About RMU Y Body Axis

MZ +7.0 Ft-Lbs Total Moment About RMU Z Body Axis

Docking Cage Spin Torque TS 0.0-5.0 Ft-Lbs 0.0-73.0 RPM Docking Cage Spin Rate +4.62 Ft-Lbs Total Jet Moment, RMU Roll Axis

+4.62 Ft-Lbs Total Jet Moment. rMU Pitch Axis +6.63 Ft-Lbs Total Jet Moment, RMU Yaw Axis

- "OR" Gate

- t"AND" Gate

S - Laplace Operator

Ws 0.0-83.0 RPM Docking Cage Inertial Angular Rate

-19- S SD 70-580 IIi l l il M j " n f ~ a 44 s. -­ -H 44-L 144U-&t

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-23- SD 70- 530 Space Division E? NorthAmencan Rockwell

TABLE 2.2.2 DEFINITIONS OF STATISTICAL DATA REDUCTION SYMBOLS

ix Arithmetic Mean

Xabs Mean of Absolute Values

S Standard Deviation

Miss Contact Radial Miss Distance (Approx. 36rvalue) between longitudinal axis of docking system and centerline of the ATS.

L = LOGe (._)whereP= .9973

SRY Standard deviation of radial miss distance in the Y axis direction

SRZ Standard deviation of radia miss distances in the Z axis direction

Contact radialvelocity (approx. 3(f VR value) between longitudinal axis of docking system and centerline of the ATS SVy Standard deviation of radial veloc­ ities in the Y axis direction

SVZ Standard deviation of radial veloc­ ities in the Z axis dtrection

-T Arithmetic mean of Y components of R23Y contact miss distance

-T Arithmetic mean of Z components of miss distance R23Z contact -T Arithmetic mean of Y components of V23 Y contact velocities

-T Arithmetic mean of Z components of V23 Z contact velocities

-24- SD 70-580 Q Space DMsion ' North Amencan Rockwell

TABLE 2.3.1 DEFINITION'OF ABBREVIATIONS

RMU Remote Maneuvering Unit

ATS-V Applications Technology Satellite, 5th of Series

FPS Feet Per Second

Spin Cage or Docking latches mounted on a symet­ rieal truss which is capable of being Docking Cage rotated about its longitudinal axis at a spin rate to match the ATS spin rate.

Transport Rig The servo driven carriage which moves the simulator camera as if it was the MT during the docking maneuver.

GPVS General purpose visual simulator includes the transport rig, a wall mounted gimbal system, and the tele­ vision systems required to simulate and visually present relative motion of one spacecraft with respect to another.

I.0. Initial conditions at the start of the maneuver in terms of relative velocities and position.

Active Vehicle EMU vehicle being flown from ground command.

Target Vehicle ATS satellite being approached for docking.

CMG's Control moment gyro control system provides attitude hold without limit cycle

DPG's Dual purpose gyros; another name for control moment gyros.

RCS Reaction control system, includes jet logic, solenoids, and reaction jets.

C.G. Center of Gravity

Center line, line of geometrical symmetry.

-25- SD 70-5bO ' Space DMsIon North American Rockwell

TABLE 2.3.1 DEFINITION OF ABBREVIATIONS (continued)

TV Television

ADIOS Analog Digital Input Output System

CSMP Continuous system modeling program, digital program which provide closed loop time histories to compare and check analog program time histories.

-26-

SD 70-580 Space Divsion NorthAmerican Rockwell

3.0 SIMULATION EQUATIONS

3.1 GENERAL DESCRIPTION

The investigation of control precision capability during renmote docking of the RMU with the ATS-V satellite, as shown in Figure 3. 1. 1, involves the simulation of vehicle relative motion in response to pilot/control system activity. The response must include a transport lag representing the time required for a signal to reach the vehicle from the ground station and the time required for the motion cues to reach the ground station from the vehicle. In the case of the RMU, remote closed circuit television furnishes the vehicle motion cues to the pilot. This requires an accurate visual simulation of the ATS-V and the RMU docking interface. The following paragraphs define the equations that must be programmed in the analog computer which will drive the television simulated RMU in response to pilot commands.

3.2 INITIAL & FINAL CONDITIONS

The general purpose visual simulator (GPVS) provides 40 ft longitudinal travel and +5 ft transverse travel without scaling. Since the transverse travel is too constraining, a 1/4 scale model will provide 160 ft. longitudin­ al and +20 ft transverse travel. Further reduction in scale impacts ability to measure miss distance accurately. The initial conditiohs of the simulation for the final phase of docking will be X = 160 ft and Zy, YT = ±10 ft. The initial range rate will be 1. 0 FPS with 3.0 degrees angular misalignment. Test conductor selection of initial conditions (see Table 3.2 . 1) will provide some variation in the pilot task. The initial condition matrix is written in terms of the RMU body axis system with respect to an inertial frame at the ATS C.G.:

TABLE 3.2. 1 INITIAL CONDITIONS

I.C. T T T T T T V14xo V14Yo Vl4Zo G T' Rl4Xo R14yo RI4 Zo Condition (FPS) (FPS) (FPS) Deg Deg Deg Ft Ft Ft

1 -1.0 0 0 0 +3 0 -160 -10 +10 2 -1.0 0 0 0 -3 0 -160 +10 - 10 3 -1.0 0 0 +3 0 0 -160 +10 +10 4 -1.0 0 0 -3 0 0 -160 -10 +10

-27- SD 70-5cC0 /

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Figure 3.1.1. RMU/ATS-V Simulation Axis System Space Divison 01 North Amerian Rockwell

The velocities, attitudes and positions when the docking interfaces make contact (R3X = 0) are the data results that indicate pilot capability for precision control to docking. These data also become docking system contact criteria and will be used as initial conditions for post contact docking loads and stability analysis. The volocities, attitudes, and positions are generated by the analog computer equations in the:axis system and form that is readily useable in the post contact docking program, that is, relative motion in the target vehicles (ATS-V) axis system. :-The data is then treated statistically to determine mean values and' standard deviations. The data reduction equations are presented in section 6.0.

3.3 AXIS SYSTEM & ROTATION ORDER

The rotation order of the RMU (active vehicle) axis system with res­ pect to the ATS-V (target vehicle) body axis system required to derive the transformation, matrixes is shown on Figure 3.3. 1. These axis systems are compatible with the body axes shown on Figure 3. 1. 1.

3.4 MATHEMATICAL MODEL

The mathematical model of the control systems, equations of motion, and TV transport drive equations are shown in the block diagram on Figure 3.4. 1. Vector-matrix notation is used in the following manner:

1) Thick arrows represent column-vector variables. 2) Thin arrows represent scalar variables. 3) Subdivided blocks represent matrix multiplications. 4) Double walled blocks represent recorded data,

Each block is identified on Figure 3.4. 1 with a number in a triangle. A functional description of each block is indexed in the fbilowing para­ graphs according to triangle number:.

(1) Rotation control rate command signals are generated by the pilot using rotation hand controller deflections. Rate commands versus deflection curves and controller stiffness are shown in Section 2.0.

(2) A control moment gyro control system (Ref. 4) is simlulated by taking pilot rate commands and converting them into control moment gyro torque commands. Momentum of the gyros is transferred to the vehicle as a moment until the commanded vehicle rate is achieved. As the gyro gimbal angle reaches 450, the desaturation switch causes the reaction jets to fire and provide control moments when the CMG's capabilities are exhausted. The CMG control system has two gyros per axis. The moment of inertia shown in the block diagram is a combined moment of inertia representing two gyros. Only one control axis is shown.

-29- SD 70-5b0 Space Division North Amencan Rockwell

XA

x

TYT XT xXA Y1

qJ Y1

Pitch (+ G About YT

_Z ! Yaw + About LzT 1 Roll ®+ About XA

FIGURE 3.3.1 ACTIVE VEHICLE ROTATION ORDER

-30-

SD 76-5o0 Alk Space Division 'K North Amencan Rockwell

(3) RCS thruster ignition signals are generated whenever the desaturation switch point of 450 is exceeded by the DPG gimbal angle. The first order lag represents solenoid actuation delays.

(4) This block represents the translation command control stick which is basically an on-off switch that commands translation acceleration. Translation stick characteristics are shown in Section 2 .0.

(5) The pulse discriminators are monostable multivibrators which assure a minimum actuation time (t min) for any signal time less than the minimum actuation time of a reaction control jet.

(6) The jet select logic determines which set of reaction jets fire in response to the various combinations of rotation and translation commands. The form shown in the block diagram is a rotation command priority system. That is, when simultaneous rotation & translation commands are introduced in some combinations, rotation jets will fire instead of translation jets. The jet nomenclatures and direction of thrust is shown on the simulation equation block diagram.

(7) Jet thrust rise ;and tail-off are simulated by first order lags to properly shape jet thrust as a function of time.

(8) The total weight of fuel consumed during each docking maneuver is determined by integrating with respect to time all twelve of the reaction jet thruster outputs. The fuel consumed during translation maneuvers is computed by integrating, with respect to time, the sum of the absolute values of the translation driving forces FX, Fy and Fz. The difference between the total fuel consumption and the translation fuel comsumption establishes the rotational fuel consumption. All three will be recorded during each run to aid in establishing the point when the pilot has "learned" the maneuver and little change in fuel consumption occurs from one run to the next.

(9) The rotational moment block is a matrix made up of moment arms about the X, Y & Z axis of the RMU for each of the 12 RCS jets. The distance G is the Y and Z dimension from the RMU centerline to each RCS "pod". The moment arm H is the one half the distance between jets along the RMU X axis.

(10) Translational forces are treated in a similar matrix in which the jet thrusts are transformed into RMU body axis force vectors by the assigned sense.

-31- SD 70-i bO Ad4k Space'Dvislon 4E3 NorthAmencan Rockwell

(11) C. G. offsets are accounted for by a matrik of X, Y & Z moment arms made up of the distance between the C.G. and the center of the reaction jet location plane. Body axis forces are multiplied by the matrix and added to the rotational moments along with dual purpose gyro moments.

(12) The rotational equations of motion shown in the block diagram have been reduced from the complete equations shown as follows: 4r=fEm1V2 +I ( WxW )+I (V+ WyW) +I z y-VW2) + w x x xWWy z2 xz ±WX Y+yz t

(Iy-I z ) WyWz + Ts]

y = y + WyWz) + Iy z (W - WxWy) + Izx \V z W ) I-1 W, w

Wz =I [Mz + Ix 4wx _ WyW z) + z (yWy +WxWz) + Ixy W2x +

(1x - Iy) WxWy + Wy J1

Note the additional terms TS, W 4 J and Wy7 J are shown in the dqua­ tions to describe the vehicle dynamics effects of docking cage spin-up torque and gyro pressession caused by the spinning docking system.

The assumptions made to reduce the equations are as follows:

(a) All multiplications of inertia times (body rate) 2 are small compared to DPG momentum torques i. e. (10) (. 01745)2 = .012<<2.0 ft lbs 2 (b) Izz Iy, Icross < 10% of Iprinc. = 4.0 slug ft ft 2 (c) IZ - Ix < 10 slug

(13) The translational equations of motion compute the translational velocity of the RMU body axis with respect to the ATS body axis including the effects of the earth orbital rate. The equations operate on the RMU body axis forces FX, F & FZ .

All W's in equations should be(j(Omega).

-32-

SD 70-580 Space Division '11 North American Rockwell

The equations shown on the block diagram were reduced from the following: F M - 24R 4 , V]14zW + 14W z S F 2 v14 y~ 0un14y 4 z x V14 XWz 14a z + 2 Q0 R 14 Z - Vl4yW x + gl4xWyVW

The equation reduction is based on the following assumptions:

(a) The orbital rate 2 : .0900727 rad/sec 2 and £ 2R = (.00007) (160) = .000011 ft/sec 2 and is<,<2 .2 ft/sec M 14,18 Both of these blocks are integrations of the incoming parameter as a function of time and are indicated by the Laplace notation'.

15,16,17These blocks are the axis system transformation matrices required to transform vectors in the RMU axis system to vectors in the ATS vehicle axis system. The trig functions have been reduced by the small angle approximation from the following matrix: (Ref. 3) cesvso C9CFCGSC ~CGsV's0 I CCGSCI +SG -sWc cws -cq1S -SG04' Ce90 CGC0 +sesqsf - sesqs

The block 15 transformation is required to provide distance signals to drive the TV camera relative to the ATS vehicle measured in the ATS axis system. The TV drive signal is generated in the ATS axis system because the ATS model is mounted on a tri-axis gimbal system and all relative attitude signals drive the gimbals. Thus the camera rig moves in relative trans­ lation degrees of freedom and the model moves in relative rotational degrees of freedom. The blocks 16 & 17 resolve relative velocity and translation data between the docking interfaces into the ATS axis system, respectively. Block 17 is the transpose of the above matrix. Data, thus generated, is in the form required for the post contact docking loads and stability analysis to be conducted on a digital post contact docking program.

*All W's in equations should be to (Omega).

-33- SD 70-580 Al Space Division 'T North Amencan Rockwell

19. Angular alignment, angular rate, linear velocity and miss distance are data time histories to be recorded during piloted simulation runs. All data is to be recorded in feet, seconds and degrees until the distance R1 5T is 1.0 ft, then switched to inches, seconds and degrees.

20. The docking cage spin motor mounted on the TV camera transport spins up a model RMU docking cage on pilot command until it is visually matched with the ATS model spin rate. The visual synchronization of spin rate is required to null relative roll motion between the RMU docking latches and the ATS docking ring. The spin rate and spin torque are also sent to the RMU rotational equations of motion such that the gyro dynamics of the spinning docking system will realistically modify the rotational control handling qualities of the RMU. The RMU roll control jets are turned on to counter torque the vehicle. The spin cage torque is balanced against the jet torgue using the roll DPG gimbal angle.

21. The model transport room contains the transport camera representing the RMU vehicle and the ATS model mounted on a gimbal system. The various points of geometry required in the simulation are identified by number. The distance between one point and another the identification numbers as subscripts. For example, R T uses 23 is the X, Y, Z distance between (2) ATS docking plane and (3) RMU docking plane J . The superscript T says the distance is measured in the target (ATS) axis system.

22. The pilot console is represented'as a blobk in the simulation diagram that receives the TV signals. In addition, it will receive angular rate and attitude information for display on the cockpit instruments described in section 4.2.

23. The signal lag block is a transport lag representing the time lag it takes for a command signal from the earth to reach the RMU in a 20,000 mile orbit and return in the form of a television view of the vehicle's response to the command. In reality there are two lags, one on command signals, and one on TV signals. However, one total lag anywhere in the loop between the command and the return picture will accomplish the same effect.

-34- SD 70-5U0 A Space Division NorthAmencan Rockwell

4.0 SPECIAL EFFECTS REQUIREMENT

The purpose of this section is to identify the special effects equip­ ment required by the RMU/ATS-V simulation.

4.1 MODEL REQUIREMENTS

Two 1/4 scale models are required for the simulation, (1) the ATS satellite shown on Figure 4.1.1, and (2) the RMU spinning docking cage shown on Figures 4.1.3 and 4.1.4.

4.1.1 ATS Model

The ATS model is configured such that the center shaft is 1.6 degrees off of the geometric center line to simulate the existing real motion of the ATS when the model center shaft is spun up to 73 RPM by the GPVS inner gimbal motor. The ATS model is also designed to permit the forward skirt to break away when inadvertent contact is made between the docking cage model and- the ATS model.

This feature will protect the GPVS Gimbal System and the expensive part of the ATS model. Since the detached skirt is very light in weight, very little damage can be done to the spinning docking cage model.

The color and reflectivity of the ATS components primarily visible during approach to docking are listed below. The components identified by circled numbers on Figure 4.1.1 are to be finished as follows:

Item

(1) Vacuum deposit aluminum HFP4-142 Type II (Skirt Rim) (2) Black Print #17038 HP4-112, Type I, Cl 2 (Aft Bulkhead and Boxes) (3) White, #17875, HP4-112, Type I, CL2 (Inside Thrust Ring) (4) Vacuum Deposit Aluminum PH4-142 Type II (Center Hub Thermal Control) (5) Vacuum Deposit Aluminum W/Black Center (Cannon Plugs) (6) Anodize MIL-A-8625 Type I (Sun Sensor) (7) Black #17038 HP4-112, Type I, CL2 (Inside Wall of Skirt) (8) Non-Gloss Black (Alternate Strips of Thermal Control) (9) Semi-Gloss Dark Blue (Equipment Section) (10) High-Gloss Dark Blue (Fwd and Aft Skirts) (11) Semi-Gloss White (Aspect Sensor) (12) Silver Mylar (Alternate Strips of Thermal Control) (13) Semi-Gloss White W/Black Center (Axial Jet)

4.1.2 RMU Spinning Docking Cage Model

The Spinning Docking Cage Model, shown on Figures 4.1.3 and 4.1.4 is not a true 1/4 scale reproduction of the RMU hardware in that the cage has

-37­ ATS

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been foreshortened and the latch triangle reduced. The distortion is re­ quired to make the docking cage model look as though it is contacting the ATS model at the end of final closure. The models would be damaged if con­ tact actually occurred.

The docking cage is mounted on the GPVS transport with the camera LOS coincident with its center line. The field of view provided by the camera is the same field as intended for the spacecraft.

The docking cage is designed to spin in roll and is driven by a servo motor to maintain 73 RPM.

Twelve small incandescent lights, color temperature 3400 0K, are attached to the camera, as shown on Figure 4.1.4 to simulate a neon circular tube light. The intensity range of the lights is variable from 1.6 to 19.2 candle power. This light ring will be filtered as necessary to simulate the spacecraft light ring.

The docking cage is painted semi-gloss white.

4.2 PILOT COMMAND CONSOLE REQUIREMENTS

The pilot console will furnish the required hardware to control and monitor the docking maneuvers. It consists of a TV monitor, six meters, four toggle and one push-button switches, two vernier potentiometers, three indictator lamps and two hand controllers.

A 17-inch TV monitor will display the ATS via closed circuit TV. The attitudes (9, 0,) which are scaled from +18 to -18 degrees will be dis­ played on three meters on the left side of the front panel. The latching index meter will display the angle offset between the RMU latches and the ATS docking ring. Readout will be +10 to -10 degrees. Two lamps, one red and one green, adjacent to the index meter will indicate when the latching index meter is in operation. Range will be displayed on a meter with 0 to 20 scaling. A pushbutton switch below the meter will enable proper range scaling selection of 0-20 feet or 0-200 feet. The rate meter will have a scaling from +1.0 to -0.5 feet per second to indicate range rate. Both range and rate meters will be located on the right side of front panel. When docking has been completed (R23 = 0), an indicator lamp on the panel will illuminate.

On the side panel, located on the right side of console, one toggle switch and a vernier potentiometer will be used to control the docking sys­ tem spin synchronization. Another switch and potentiometer will control visual synchronization. A third switch will select either the spinning image or the despun image to be displayed on the T.V. monitor.

Two hand controllers will be mounted in front of the console. The rotational controller will be for right hand operation and the translational controller will be mounted for the left hand to operate.

-42- SD 70-58o D7PILOTf. MaVJeTort

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I-IQRE4.z-/ PILOT- CONTR.-OL- COiJStLE­ dihk Space Division NorthAmerican Rockwell

4.3 VISUAL SYSTEM REQUIREMENTS

The visual displays for the RMU/ATS docking simulation shall consist of a 525 line television display. The system shall consist of three elements: (1) RMU TV camera, (2) derotation optics display system, and (3) pilot con­ sole display.

4.3.1 System Description

A 525 line 'TV camera shall be mounted on the General Purpose Visual System (GPVS) transport to represent the RMU on-board camera. The image from this camera shall be displayed on a TV monitor (Figure 4.3.1) and reviewed through the derotation optics system with a second TV camera. The image from the second camera shall be displayed on a monitor in the pilot console. A switch shall be provided to allow the image from either TV camera to be displayed on the pilot console monitor. If RMU attitude exceeds +6' in pitch or roll the video-signals to the pilots monitor will be interrupted to simulate loss of RMU antenna alignment.

4.3.2 TV Camera

The RMU TV camera shall be a COHU Model 2000 525 line camera. This camera is a 10 mhz system. A 10 mm lens shall be used to provide a 70' field of view.

4.3.3 Derotation Optics System

The derotation optics TV monitor shall be a nine inch 525 line Conrac TV monitor. The derotation optics prism (Reference 5) shall be positioned and aligned between the monitor and a second COHU Model 2000 TV camera. The derotation optics shall consist of a servo-driven DOVE prism (Figure 4.3.2). Since the derotation optics gives a mirror image, the inversion will be corrected by reversing the TV image electronically.

4.3.4 Pilot Display

The pilot console shall have a 17 inch 525 line TV monitor which dis­ plays either the direct RNU video or the derotation optics video. The television system shall provide at least 350 lines vertical resolution and 500 lines horizontal resolution.

SD 70-580 RMU CAMERA AT5 MODEL

tS & 5 AT-OR-

it II TArc IO OPTICS GROUND

tt v

a_- S, J-rt l /CKA'j:tt 0 0 W OtJI TCR

FIc*L i 4L-73.! V,'h'citu, 0. \' t,'. . 'C Figure 4.3.2 Derotat n Optics System SD 70-5 A Space Division ' North American Rockwell

5.0 PROBLEM CONTROL

The purpose of this section is to specify the basic areas of responsibility of the participating groups, the requirements of the simulator, and the control of the test.

5.1 RESPONSIBILITIES

S.1.1 Sponsoring Department

The sponsoring department (Advanced Programs D/254) will provide definition of study objectives, parameters to be-investigated, simulator design requirements, equations of motion, test plan and run schedule, and monitoring of the simulation to insure that the study objectives are achieved. .Pilot debriefing, final reduction of data, and writing the simulation data report will be accomplished by the monitor. Mr. Glynn Mount, Jr. (D/191) has been assigned this task.

5.1.2 Simulation Project Engineer

The simulation project engineer will provide and negotiate the support work authorization (SWA) for simulation support, procure simulation hardware, provide the direction of all simulation personnel to assure that the simulation meets all accuracy requirements throughout the evaluation, main­ tain data records and configuration control, and assist in writing the final data report. All decisions affecting the scheduling and objectives of the study will be reviewed with the sponsor group. Mr. Don Wurzburg (D/190) has been assigned this task.

5.1.3 Simulation Personnel

Simulation personnel will scale and mechanize the simulation equations, design and install simulation models and control hardware, conduct pre­ simulatipn software and hardware checkout and verification, conduct daily verification of the simulator, monitor, log, and submit all data related to the study to the data files at the end of each day.

GPVS Engineer - Mr. Norm Svensrud is responsible for the operation of the General Purpose Visual Simulator (GPVS), the simulation models, the pilot console and their interface with the analog computer.

Analog Computer Engineer - Mr. Bob Morris is responsible for the scaling and mechanization of the simulation equations and control functions on the analog computer.

Analog Operator - Mr. Chuck Morally is responsible for the maintenance and operation of the analog computer and data recorders through­ out the study and provides programming assistance to the computer engineer.

SDS 70-047- 58o QSpace Division 9 NorthArencan Rockwell

Simulation Technician - Mr. Rich Williams is responsible for the installation and maintenance of all simulation hardware and wiring inter­ faces between the computer, pilot console, GPVS, models and visual system.

5.1.4 Pilot Subjects

There will be a minimum of three NR test pilots assigned on a continuous standby basis. The same pilots must be available throughout the entire production period. This avoids the necessity of training additional pilots and prevents data scatter in a relatively small number of run replications from biasing statistical results. The pilot's task will be to develop the docking maneuver procedures, recommend control system sensitivities based on an accepted grading method, suggest improvements in any of the vehicle systems that may improve their ability to approach docking contact with extreme precision.

-48- SD 70-580 Adik Space Division 'FT North American Rockwell

5.2 SIMULATOR ACCURACIES

5.2.1 General Purpose Visual Display System

The three-DOF transport of the visual display system consists of a servo-driven carriage system which can position the TV camera. This assembly is used to simulate both translation and rotation of the active vehicle. The transport consists of three independently driven orthogonal carriages. The performance characteristics of the transport are given in table 5.2. 1. This table, lists two sets of characteristics, (1) the overall system capability, and (2) a limited travel capability. The latter capability describes the performance at docking contact for final conditions.

TABLE 5.2.1 RIG PERFORMANCE

X Y Z Characteristics Overall Limited Overall Limited Overall Limited

Position Accuracy +0.4 in. +0.10 in. +0.1 in. +0.05 in. +0.05 in. +0.02 in.

Displacement Max. 43 ft - 9 ft. - 17 ft. -

Velocity Max. 2.5 ft/sec .8 ft/sec - .3 ft/ sec. - 3 - 3 Velocity Min. .15x10- 3 .15x10 - .15x10 ft/sec ft/sec ft/sec

The characteristics stated in table 5.2. 1 are the unscaled GPVS characteristics and must be multiplied by 4 to reflect the quarter scale model characteristics.

A three-DOF gimbal system is used to mount the model of the ATS. This gimbal system is used to provide both the spin rate of the ATS and maintain the proper relative angular alignment between the RMU and ATS. The characteristics of this gimbal system are indicated in table 5.2.2.

-49- SD 70-580 Ik Space Division North American Rocwell

TABLE 5. 2.2 GIMBAL PERFORMANCE

Characteristics Inner Gimbal Middle Gimbal Outer Gimbal

Position Accuracy +.250 .250 .250

Maximum Rate 480 0/sec 40 0 /sec 40 0 /sec

Range Continuous ±1000 -+ 1000

5.2.2 Television Monitoring System

The television monitoring system consists of a camera mounted on the transport; the view from this camera may be sent directly to the pilot's console or routed through de-rotation optics. The derotation optics consist of a dove prism which is rotated at 1/2 the speed of the ATS. An image, presented on an alternate monitor, is viewed through the dove prism by a second television camera and then routed to the pilot's console. The overall TV system will be 500 lines at the pilot's monitor.

Alignment of the dove prism, monitor and camera is accurate to within .01 degrees.

5.2.3 Displays

Exclusive of the television monitor,the pilot displays consist of vehicle attitude relative to the ATS, range, range rate and relative attitude between the RMU Docking Cage and the ATS, range, range rate and relative attitude between the RMU Docking Cage and the ATS thrust ring as indicated in figure 4.2.1. These, displays are provided by vertical scale meters. Vehicle attitude will range from +180 to -180 plus or minus 0.250 in all axes. Range Rate varies from -0.5 to 1.0 feet per second and can be read to plus or minus 0. 1 feet/second with an accuracy of .05 feet per second. Two scales are provided for the range indicator, the first 200-0 feet in increments of 20 feet and the second 20-0 in increments of 2 feet, On the fine scale the range indication is accurate to within 3 inches. The latch index display provides relative attitude between the ATS and the RMU docking cage. This indicator is only operational after the relative angular spin rate has decreased below 3 RPM and the relative attitude is within + 10 degrees. The indicator is calibrated in 1.0 degree increments and is accurate to + .025 degrees.

-50- SD 70-58o Space DiMsion Northkerican Rockwell

5.2.4 Controls

The docking cage and derotation optics controls consist of a fixed initial command of 73 RPM and a vernier adjustment from the pilot's console.

The dove prism may be accelerated to 73 RPM in 10 seconds by the fixed command. The docking cage is accelerated to 73 RPM in 60 seconds by fixed command. Both servo systems are capable of holding the set point within 0. 1 percent and provide vernier control of +3 RPM.

The rotational hand controller characteristics and tolerances are presented in figures 2.1.2 through 2.1.5.

-51- S SD 70-580 Space Division 90 NorthAmencan Rockwell

5.3 SIMULATION CHECKOUT REQUIREMENTS

Validation of the RMU-ATS simulation will be composed of an initial system verificatio- and periodic revalidation. The initial system verifi­ cation will provide for open loop checkout of the individual system performance parts and closed loop dynamic checkout of the integrated system. The periodic revalidation will consist of a daily checkout procedures with more sophisticated procedures as required by changes to the initial baseline.

5.3.1 Open Loop Checkout

5.3.1.1 Pilot Console

a) Determine the rotational controller breakout deadbands for roll, pitch and yaw.

b)' Plot command voltage vs. stick deflection of the rotational controller for roll, pitch and yaw.

c) Determine the translational controller deadbands in X, Y, and Z.

d) Verify the relative attitude indicators for relative attitudes of +5, 10, and 15 degrees in roll, pitch and yaw. These readings shall be accurate to 0.5 degrees.

e) Verify the range rate indicator from -0.5 to 1 foot per second in steps of .25 feet per second.

f) Verify the range display in both scales (0-200 and 0-20) for 5 equally spaced points over the range. This display shall be accurate to 10%.

g) Verify the ATS/Docking cage relative attitude indicator over the range of +10 degrees. - This display shall be operational only when the relative rates are below 3 RPM, and the relative attitudes are less than 10 degrees. For rates and attitudes outside these bands the red indicator shall be on. In the operational region of the indicator a green indicator shall be on.

h) Determine the spin up time of the docking cage from the initial command in seconds.

i) Determine the drift rate of the docking cage (.1%).

j) Verify that the venier control has +3 RPM authority around the 73 RPM nominal.

k) Determine that the spin up time of the dove prism from the initial command is 10 + 2 seconds.

1) Determine the drift rate of the dove prism.

-52- SD 70-580 9dl Space Divsion North American Rockwell

m) Verify that the dove prism venier control has +1.5 RPM authority around the 36.5 RPM nominal.

5.3.1.2 Television System

a) Verify picture resolution on both the spinning and de-spun images.

b) Verify proper lighting of the ATS model.

c) Verify loss of video signal with pitch or roll angles in excess of 6 degrees.

d) Determine dove prism misalignment with the monitor and relay.

e) Adjust the. RMU range to 50 feet and verify the television monitor 50 foot range circle coincides with the ATS image. Repeat for a range of 5 feet.

5.3.1.3 General Purpose Visual Displays

a) Verify alignment of the ATS and RMU reference'frames. Bias to align body axes to within +.25'.

b) Verify relative motion of the transport and gimbal system for rotational and translational commands.

c) Verify the 1.60 offset of ATS model.

d) Verify that the transport moves .25, 1.25, and 2.5 feet respectively for translation commands of 1, 5, and 10 feet in X, Y, and Z.

e) Determine the transport position for the RNU rotational maneuvers listed in-Table 5.3.1.3a.

Table 5.3.1.3a RNU Checkout Maneuvers (Position Check)

Range Rotational Command Transport

______QC ______XY F- Z +3 +2.09' 160' +3 +2.09' +3 -- +3 +2.09' +2.09'

+3 +.209' 16' +3 +.209' 73 +3 +.209' +.209'

-53- SD 70-580 Ik Space Division, NorthAmerican Rockwell

f) Establish the transport rates for the RMU rotational maneuvers

as listed in Table 5.3.1. 3b.

Table 5.3.l.3b. RMU Checkout Maneuvers (Velocity Check)

Range Rotational Command Transport ''0y z

50 ft. . deg/sec I deg/sec 10.46 in/sec 10.46 in/sec

8 ft. 1deg/sec I deg/sec 1.67 in/sec 1.67 in/sec

Verify that the ATS gimbal rates reflect the appropriate RMU

rptational rates.

5.3.1.4 ATS Model

a) Determine the wobble amplitude of the ATS docking ring. The

amplitude of the wobble perpendicular to the spin axis should be

0. 568 inch, while that parallel to the axis is 0. 5Z0 inch.

b) Verify that the spin rate of the ATS model is 73 + . I rpm.

5.3. 1.5 Control System

a) Verify that the time lag of the rotational controller command

signals is 220 * 5 milliseconds.

b) Supply torque command of + .5, + .75, + 1.0 and + L.5 deg/sec

to the CMG torquer and measure the torquer output. The corres­

ponding torquer outputs are - . 00873, + .02616, • .03490, and

- .03490 ft/lbs.

-54- SD 70-580 AI~l Space Division NorthAmencan Rockwell

c) Supply a torque command of 1. 0 deg/sec to the CMG torquer with

the rate of feedback disabled. Measure a, a , and MDPG as a

function of time and determine a when the desaturate switch

activates. Compare with digital results.

d) Supply a step input of 45 degrees to the control lag circuits and

record the output. Measure the time constants (10 + I milli­

seconds).

e) Supply a five millisecond command to each of the one-shot milli­

vibrators and record the output. Verify the output pulse width

is 12 + 1 millisecond.

f) Verify that the time lag of the rotational controller signals is

220 -t 5 milliseconds.

g) Verify the translational controller inputs according to Table

5.3.1.5a.

Table 5.3.1. 5a. RMU Jet Checkout (Translation)

Controller Input Reaction Jet

-x C3D3 + x A3 B3 +Y AZDZ -y BZCZ -Z BIDI + X AlCl

-55- SD 70-580 491h SPa c e vDI sIon North American Rockwell

h) Verify rotational control prioiity. Command translational and

rotational comnands as indicated in Table 5.3.1. 5b and verify

that the indicated jets are inhibited. Disable the vehicle rate

feedback. The inhibited jets will turn off when the DPG's reach

+ 45 degrees switching limit.

Table 5.3.1. 5b. RMU Jet Select Checkout

Translational Rotational Translational Rotational Final Jets Control Control Jets Jets

- x +0 C3 D3 C3 B3 C3 B3 - x - Q C3D 3 A3D 3 A3 D3

- x - CD 3 A 2 C2 C3D3A C 2

- x + F C3 D3 B2 D2 C3 D3 B2 D2

- X - 0 C3D3 AIB 1 C3 D3 AIB1

- x + 0 C3 D3 CID 1 C3 D3 CID1

+ X + 0 A 3 B3 G3 B3 C3 B3

+ X -Q A3 B3 A 3 D3 A 3 D3

+ X + A 3 B3 A 2 C2 A 3 B 3AZC 2

+ X -W A 3 B 3 BD 2 A 3 B 3 BZD2

+ X +0 A3 B3 AIB 1 A 3 B3 A1 31

+ X - 0 A 3 B3 CID 1 A3B 3 CDl1

+ Y +0 AZD 2 C3 B 3 AZD 2 C 3 B3

+ Y - 0 A 2D2 A 3 D 3 AZD 2A 3D 3

-56- SD 70-58o 0% SpaceDivsion North American Rockwell

Table 5.3.1. 5b (Cont'd) RMU Jet Select Checkout

Translational Rotational Translational Rotational Control Control Jets Jets Final Jets

+Y- AZD' A 2 C2 A 2 C Z

+Y + IF A2 D 2 B2 D2 BzD 2

+ Y - 0 A 2 D2 A1 BI A 2 D2 A3I1

+ Y + 0 A2 Dz 2 CD l A2 D2 CD l

- y + Q B 2 C z C3 B 3 BzC2 C3 B 3

- Y - 0 B2 C2 A 3 D 3 BzC2 A 3 D 3

- y -Bl B2 C2 A 2 C 2 A 2 C2

- y +q B 2 C 2 BzD2 B2

- Y - 0 B 2 C 2 AIB1 B2 C 2 AIB

- y +0 B 2 C2 C1 D 1 BzCzCIDI

- Z + 0 BID l C 3 B 3 BD 1CB 3

- Z - 0 B1 D1 A 3 D 3 BIDiA 3 D 3

- Z -T B D1 A2 C2 BID1A 2 C2

- Z +WI BID 1 B 2 D2 BlDIB2D Z

- z -0 BIDI AlB 1 AB I

- z + 0 BID 1 CID 1 CID,

+ Z + 0 A C1 C3B 3 AlIG3B 3

+ Z - 0 A 1C1 A 3 D3 AICIA 3 D 3

-57- SD 70-580 Space Division .1NozthAmerican Rockwell

Table 5.3. 1. 5b (Cont'd) RMU Jet Select Checkout

Translational Rotational Translational Rotational Final Jets Control Control Jets Jets

SZ -T AIC A A 0IIAZC Z

+ Z +4! A1 Cl BZD 2 A1 GIBZD Z

" Z -0 A1 C AAB

" Z + 0 AIC1 CID I CID 1

i) Verify that the rise and fall times of the reaction jets is 38milli­

seconds and that the thrust level is L. 5 - . 01 pounds.

-58- SD 70-580 Space Division North American Rockwell

5.3.2. Closed Loop Dynamic Checkout

The purpose of the closed loop dynamic verification is to ensure that when the components of the simulation system are combined in a closed loop manner, the response of the simulated system represents the RMU system's ex­ pected performance as predicted by analysis.

5.3.2.1 Analog Math Model

Closed loop checkout of the analog mechanization will be accomplished by comparison with a set of single axis digital runs. The checkout run schedule will contain sufficient run types and initial conditions to verify the closed loop response of the control system and vehicle dynamics.

5.3.2.2 Visual Verification

Fixed translational and rotational commands will be supplied to the system for predetermined periods of time; correlation will be established between the Television System, Analog Computers, displays, and GPVS Trans­ port.

5.3.2.3 Closed Loop Inhibit

Verification of the rotational priority system will be established by simultaneously subjecting the system to rotational and translational commands. The system should display both translation and rotation maneuvers until the DPG's saturate at which time the translational rates become constant. Record vehicle angular and translational rates, DPG angle a, and reaction jet thrust.

5.3.3 Continuous System Verification

The continuous system verification will consist of morning and evening static and dynamic checks as required to maintain an accurate account of the system configuration. These checks will consist of pilot console verifi­ cation, analog computer static and dynamic checks; visual display alignment and scaling and interface verification.

5.3.4 Revalidation

These tests will be conducted to reverify the system subsequent to system failure, or to revalidate the system and establish a new baseline, in the event of changes to the system configuration.

The tests will consist of applicable sections of the open and closed loop system verification tests as presented above.

-59- SD 70-580 SSwe DMslan . NorthAmerican Rockwell

5.4 PRELIMINARY PILOTING PROCEDURES

After terminalrendezvous and visual acquisition of the ATS-V, the RMU is maneuvered into the station keeping mode at one hundred to fifty feet from the ATS with a relative closing velocity of approximately 1.0 FPS. Attitude maneuvers at this stage should not be attempted since large maneuvers (+60) will cause loss of visual acquisition by moving the directional antenna off its earth based receivers. The maneuver from station-keeping to docking contact involves the following steps:

1. Deceleration and Line-up Translation 2. Fine Attitude Alignment 3. Translation to Center 4. De-spin Visual Image 5. Spin Up Docking Cage and Sync. 6. Index Latches 7. Clbse to Contact on Center

These steps are expanded to provide insight into the docking task as well as to suggest a standardized pilot procedure and use of pilot console instrumen­ tation.

1. Deceleration and Line-up Translation - The initial condition of each run will locate the RMU 160 ft from the ATS with a radial offset of 10 ft and a small attitude misalignment. The closing velocity will be 1.0 FPS. A continuous retro-thrust for approximately five seconds will reduce the closing velocity to zero in approximately 2.5 feet.

The suggested procedure is to step the closing velocity down to .5 FPS at approximately fifty feet as indicated either by the range and range rate meters or by the back-up concentric ring overlays on the TV monitor. Align in lateral translation and further dece­ lerate to zero closing rate at ten feet before attempting angular alignment.

Note: Angular alignment is delayed until the RMU is in near vicinity of the ATS for three reasons: (1) the pilot will not be able to see the satellite well enough at 160 feet, (2) scaling the problem to fit the TV rig room, and available degrees of freedom causes attitude changes to demand lateral translation in relative motion which may exceed permitted rig travel. If immediate attitude commands are desired, the initial condition of 160 feet may be changed to fifty feet without impairing pilot tasks, and (3) attitude maneuvers in excess of ground link antenna lobe width of 120 (+60) will cause loss of visual image.

-6o- SD 70-580 A& Space Division 9 North Amencan Rockwell

2. Fine Attitude Alignment - While maintaining a distance of between five and ten feet between the RMU and ATS as indicated by another concentric overlay on the monitor image, maneuver in either +Y or +Z translation to sight along the edges of the ATS and correct attitude alignment until the lighted edge of the ATS just disappears. The baseline rotational control sensitivity is 1.0 deg/sec vehicle command per degree of controller deflection. Translate to. the opposite edge of the ATS to make any further attitude correction in that axis. Translate to another edge 900 to the last and make similar corrections to attitude alignment. Check to see if fine alignment is verified by relative attitude meters. Correct roll attitude if required.

Note: The roll attitude meter will display a roll attitude error with respect to the optimum antenna direction instead of with res­ pect to the ATS. The pitch and yaw meters will display an attitude error with respect to a non-spinning or inertial axis of the ATS.

3. Translation to Center - Once attitude alignment has been achieved the DPG control system should maintain a tight attitude hold with little attention. From this point on, the primary task is one of translating the RMU center line to the center of the ATS using either concentric ring overlays of the monitor or the visible por­ tion of the docking system.

4. De-spin Visual Image - A switch is provided on the pilot console which spins up a ground based image derotational device. A fine vernier control is provided that permits fine adjustment of the visual despin device such that the image roll rate is visually nulled. If the image was not centered, the despin optics will cause the image to traverse the monitor around the center of the screen. To determine the translation direction required to center the ATS image it is necessary to switch back to the spinning image. This can be done,without turning off the despin optics,by activa­ ting the "spin/despin view" switch.

5. Spin up Docking Cage and Synchronize - The RMU docking system spinup is activated with a switch and a vernier control similar to the despin optics except that vehicle motion is generated by the torque applied to the docking system. The spinup torque and its duration have been chosen to provide as little disturbance to the vehicle as possible. A waiting period of two to three seconds is required for the spinup to reach a constant rate. During this period the DPG's and RMU jets will try to cancel any RMU roll rate. Once the docking system is synchronized with the ATS spin rate, some correction of the RMU roll attitude may Le necessary. Simultaneously, the pilot is required to keep the vehicles centered in radial translation.

-61- SD 70-580 . Space Division NorthAmerican Rockwell

6. Index Latches - The RMU docking latches will have zero relative roll rate with respect to the ATS if.properly synchronized. The pilot has a view of the despun image of the RMU docking system and the ATS docking interface. A minimum requirement is to index the latches around the cannon plugs visible on the ATS. A latch index indicator has been provided on the pilot console, which, if centered on zero, will index the latch around the cannon plugs as well as black boxes on the ATS aft bulkhead. The movement of the latches with respect to the ATS is controlled by the spin cage vernier on the pilot console.

7. Close to Contact on Center - The latches should be indexed, the relative attitudes zeroed, and the vehicle centerlines lined up. The remaining task is to close so that the latches are triggered by the ATS thrust ring. The distance between the RMU and the ATS should be five to ten feet. A one second pulse on the translation controller will set up a closing velocity of .2 FPS. At a distance between RMU and ATS representing contact, a latch trigger light on the pilot console indicates the end of the maneuver.

5.5 TEST PLAN AND SCHEDULE

The test plan ground rules are as follows:

1. A minimum of three NR test pilots per parameter variation. 2. A minimum of ten runs per pilot on each parameter variation. 3. Time schedule is based on an eight hour day with 50% down time for daily verification, trouble shooting, pilot debriefing and Setting up initial conditions. 4. The scope of the piloted runs includes the following:

a) Learning runs using baseline systems b) Translation control power (2 variations) c) Rotational control sensitivity (2 variations) d) Final runs on the modified baseline. 5. Evaluation of visual presentation is continuous throughout the piloted runs without devoting special ruxis for this purpose. These include lighting intensities, TV image, concentric circle overlays versus range and range rate meters, relative attitude meters and latch indexing meter versus TV presentation. 6. The initial conditions will consist of four combinations of lateral displacement, i.e. +10 feet in Y and +10 feet in Z with three degree misalignments. The initial range rate will be 160 feet and 1.0 FPS respectively. Each of these four initial condi­ tions will be flown three times per iilot per variation.

-62- SD 70-580 Adikk Space Division 9F North Amencan Rockwell

7. The time to make one run is estimated as follows: a) Closure:

Drift 110' @ 1 FPS = 110 see Decelerate to .5 FPS = 2 Drift 4o' @ .5 FPS = 8o

Decelerate to 0 FPS - 2

Time 194 sec 194 see b) Angular Alignment:

Translate top-5' @ .1 FPS = 50 sec Align = 30 Translate bottom - 50 Align = 30 Translate left 5' @ .1 50 Align = 30 Translate right = 50

Align - 30 320 see 320 sec

c) Spin Up Docking System:

Spin up 73 RPM = 60 Sync. Adjust. = 20 TU-ec 80 sec

d) Translate to Center: 3' @ .1FPS = 30 Align = 30 60 see 60 sec

e) Closure: 5' @ .1 FPS 50 sec 50 seC Total Time 744 sec

12 min. to fly 3 min. get ready 15-min . I hr. per run

S763- SD 70- 580 Aft Space Division North American Rockwell

8. Total run time per evaluation is based on the following equation:

Run time = (No. conditions) (No. times through) (No. variations) (No. pilots) (No. hrs/run)

Thus, Learning run time = 9 hours Translation Control Evaluation 18 hours Rotation Control Evaluation 18 hours Modified Baseline Run Time 9 hours

Total Piloted Run Time = 54 hours

The test schedule shown on Figure 5.5.1 was developed from the above ground rules.

The test plan is shown on Table 5.5.1. The first four columns will make up the run number for data records. For example, the first run would be identified by OllAlOO1, the 12th run would be 124A1002.

The tentative daily schedule is as follows:

8-9 AM Morning System Checkout 9-10 AM Contingency 10-12 AM Pilot Runs and Debriefings 12-1 PM Lunch 1-3 PM Pilot Runs and Debriefings 3-4 PM Contingency 4-5 PM System Verification

-64- SD 70-58o WEEKS FROM TOTAL SYSTEM CHECKOUT

Nov. 1 Nov. 8 Nov. 15 Nov. 22 Nov. 29

Evaluation Runs 1 2.. 3 4

Learning

Translation Control Pwr. m

Rotation Control Sens. Baseline I

Data Report

to co 0"0z c/ Figure 5.5.1 RMU/ATS DOCKING SIMULATION SCHEDULE 3 C

ni:I <6

CD TABLE 5.5.1 RMU/ATS DOCKING TEST PLAN

RUN COND. PILOT DATE VARIATION REMARKS NO. NO. I

1 1 A 10/1/70 Learning Runs 2 2 3 3 4 4 5 1 6 2 7 3 8 4 9 1 10 2 11 3 ' 12 4 A 10/2/70 Learning Runs oN 13 1 A Translation Control 14 2 Power Halved 15 3 16 4 17 1 18 2 - 19 3 I 20 4 a 21 1 ZO. 22 2 0 23 3 Translation Control B 24 4 A Power Halved <

0CD RMU/ATS DOCKING TEST PLAN

RUN COND. PILOT DATE VARIATION REMARKS No. NO.

25 1 A Translation Control 26 2 Power Doubled 27 3 28 4 29 1 30 2 31 3 32 4 33 1 34 2 35 3 Translation Control 36 4 A Power Doubled

37 1 A Rotation Control 38 2 Sensitivity 39 3 _c = .2 Deg/Sec. 40 4 6s Deg. W 41 1 42 2 o 43 3 0 44 4 o 45 1 .01zw) 46 2 Rotation Control .0t, 47 3 Sensitivity > m 48 4 A G=c = .2 Deg/Sec. s Deg. 0 C) o0 RM/ATS DOCKING TEST PLA

RUN COND. PILOT DATE VARIATION REMARKS NO. NO.

49 1 A Rotation Control 50 2 Sensitivity 51 3 0 = .(5 Deg/Sec. 52 4 feg. D- 53 1 54 2 55 3 56 4 57 1 58 2 Rotation Control 59 3 Sensitivity 60 4 A _ = .05 Deg/Sec. s Deg. 61 1 A Modified Baseline OCE 62 2 Statistical Runs 63 3 64 4 65 1 66 2 67 3 68 4 69 1 70 2 71 3 A Modified Baseline Statistical Runs o Zen 0 10 0

0

0 9­

(D. ' Space Division North American Rockwell

6.0 DATA ACQUISITION & REDUCTION

6.1 OUTPUT DATA REQUIREMENTS

The simulation output data will consist of three forms; (1) analog strip chart time histories, (2) ADIOS typewriter output of finial conditions, (3) pilot debriefing. Table 6. 1. 1 presents the parameter to be recorded, maximum magnitude, magnitude at the end of the run, and the units. The analog and ADIOS data will consist of angular alignment, angular rate, linear velocity, miss distance, and fuel accounting.

6.2 DATA REDUCTION EQUATIONS

The parameters marked by an asterisk in Table 6. 1. 1 will be statistically reduced for the final simulation data report. Table 6.2.1 is a typical presentation of the final form of the docking data. The data presented in Table 6.2.1 represent the means of the absolute values of all parameters except angular misalignment. Limit values represent mean + 3 d' or the highest value to occur in 99.74 percent of all probable cases. The limit value will not be used for closing velocity but rather the'actual value of the worst case. This statistic is chosen due to the highly skewed nature of the distribution as experienced in previous docking simulation studies. The statistical formulation for limit value does not apply readily to such a distribution.

6.3 DATA RECORDS

A complete set of data records of all phases of the simulation will be maintained. These records will consist primarily of Laboratory Data Books, Analog strip charts, and ADIOS printouts. A daily production run log will be maintained reflecting the status of the complex, the production run number of each run attempted, the pilot's name, initial run conditions, and comments. A daily Checkout Log will be maintained containing the results of the morning and evening checkouts.

-69- Sn 70-580 Ad Space Division 9 North Amencan Rockwell

6.3.1 CHECKOUT RECORDS

A formal set of checkout records will be established containing the results of both open and closed loop complex verification. Open loop and subsystem tests will be recorded in Laboratory Data Books where applicable, closed loop test results, consisting primarily of analog strip charts, will be recorded in loose-leaf notebooks together with the correspon­ ding digital results. Subsequent to the completion of the complex checkout a Simulation checkout report will be compiled. This document will contain a brief description of each test (open and closed loop), the expected results and the test results.

6.3.2 PRODUCTION RECORDS

All production runs regardless of degree of success will be documented. Each production run will contain a run number directly correlatable to a given set of conditions as described in the Run Schedule. The strip chart recordings will be maintained in original form and reproductions used as required for reports.

All production runs will be identified by a run number and all record­ ed parameters will be marked in the appropriate units.

6.4 QUICK-LOOK DOCUMENTATION

Since the number of data parameters is extensive,' and the oscilo­ graph recordings will be fairly long, a quick-look Data sheet will be used to summarize each piloted run. Figure 6.4. 1 presents the quick-look data sheet format. Final analysis of the quick-look data will generate means, maximums, and standard deviations of each of the parameters shown.

-70- SD 70-580 11 Space Division *41 NorthAmerican Rockwell

FIGURE 6.4.1 QUICK-LOOK DATA SHEET

RUN NO. PILOT

DE-BRIEFING COMMENTS: (Print)

FINAL CONTACT CONDITIONS:

ZMISS = _ _ inches X = IPS

YMISS = inches Y = __ IPS

G = Deg. Z = IPS

9f = Deg. G = Deg/sec

Vt = _Deg. ( = _ Deg/sec

= _Deg/sec

FUEL USED:

TOTAL LBS, ROTATIONAL LBS, TRANS LBS.

-71- SD 70-580 Miss Transverse Closing Attitude Angular Propellant Mission No. Distance Velocity Velocity Control Rate Weight Time of (ft) (ft/see) (ft/sec) (deg) (deg) (ib) (ib) Runs

a Mean Limit Mean Limit Mean Limit Mean Limit Mean Limit Mean Limit Mean Limit CONDITION

Nominal

Sensitivity Value No. 1

R) Sensitivity Value No. 2

0 o TABLE 6.2.1 RMU DOCKING SIMULATION RESULTS zw

>

CD. 9 Space Division North American Rockwell

The statistical data reduction equations axe presented as follows:

Arithmetic means: N Xi Where N = number of samples =- Xi = individual sample value 1.-i

Means of Absolute Values; N -- 1Ix I XABS =3 N 1=l

Standard Deviation;

N 2 s (xi- ) N-1 i=1

Select Maximum of Absolute Values;

Nax

Mean + 3c ;

Where K = 3 for 99.74% of total Population

For each case the following two variables are computed:

Miss Distance 99.74% values;

MIcSS = A (M + (S +

Where L = 1.725

Radial velocity 99.74% values: 2 VR - (v,2 2. (721Z) + L (s + svz) R '~ \J~d -(See Sec. 2.0 for definitions)

-73- SD 70-580 Alk S pac e D M s io n 9r NorthAmerican Rockwell

TABLE 6.1.1 RECORDED DATA

Parameter Definition of Max End Run Recorder Symbol Parameter Units Range Magnitude No.

RT23Y Miss Dist. Y Ft. +10.0 1.0 inch 1 RT23z* Miss Dist. Z Ft. +10.0 1.0 inch 1

RT23X Contact Pt. Range Ft. +160. 0.0 1

9 * Relative Pitch Degrees +10. +3.0 Deg 1 0 Inertial Roll Degrees +10. +3.0 Deg 1 * Relative Yaw Degrees +10. +3.0 Deg 1

VT4X* Closing Velocity FPS +1.0 -.5 .1 2 VT4Y* Y Rel. Vel. FPS +.5 .I 2 VT4Z* Z Rel. Vl. FPS +.5 .i 2

X * Inertial Roll Rate Deg/Sec +1.0 DPS +.05 DPS 2

WY * Inertial Pitch Rate Deg/Sec +1.0 DPS +.05 DPS 2

Z * Inertial Yaw Rate Deg/Sec +1.0 DPS +.05 DPS 2

WTRANS Translation Fuel LBS 5 LBS 5 LBS 3 WROT Rotation Fuel LBS 5 LBS 5 LBS 3

WTOT * Total Fuel LBS 5 LBS 5 LBS 3

t * Time Sec 900 900 3

O DPG Gimbal Angles Degrees +50 -- 3 TO DPG Gimbal Angles Degrees +50 -- 3 DPG Gimbal Angles Degrees +50 3

* To be statistically reduced

-74- SD 70-580 Ad Space Division 9&North Amencan Rockwell

REFERENCES

1. SD 70-370, "Proposal for a Study of Automated Rendezvous and Docking for ATS-V Despin", 3 August 1970.

*2. IL 697-523-110-64-207, "Docking Study "C" Simulation Requirements", (Apollo) 30 April 1964.

3. SID 65-1052, "Detailed Test Plan Docking System Development Test Program (MSC), Volume III - Simulation Equations", 14 February 1966.

4. IL 190-300-G&N-68-76, "Effect of Control Moment Gyro Dynamic Load Liit tions on Maeuvering Unit Performance", July 19, 1968.

5. MIL-HDBK-141, Mirror and Prism Systems

* Included for Reference Reading

-75- SD 70-580 Space Division NorthAmencan Rockwell

APPENDIX C

PILOTED DOCKING SIMULATION STUDY

SD 71-286 J91 Space Division IVV North American Rockwell

TABLE OF CONTENTS

Section Page

1.0 DISCUSSION OF THE RMU DOCKING TASK...... 1

2.0 SIMULATION EQUIPMENT DESCRIPTION...... 3

3.0 PILOTING PROCEDURES ...... 13

4.0 STUDY PARAMETERS ...... 17

5.0 SIMULATION MECHANIZATION ...... 19

6.0 STUDY RESULTS ...... 27

7.0 DISCUSSION OF RESULTS ...... 41

8.0 PILOT DATA ...... 49

9.0 CONCLUSIONS AND RECOMMENDATIONS ...... 59

PRE 4DbG PAGE .LALK NOT FIaJ

C-hi SD 71-286 Space Division E1j North American Rockwell

1.0 DISCUSSION OF THE RMU DOCKING TASK

The ATS-V satellite is presently in a stationary orbit (;-,20, 000 nautical miles) spinning in a stable condition at 73 RPM about its longitudinal axis. The spin is undesirable because it prevents deployment of ATS experiments. The RMU is to rendez­ vous, dock, rigidize and despin the ATS-V by reacting to control commands from the ground.

-The RMU piloted simulator was designed to provide data on the precision of control provided by a ground-based pilot using a television visual system as his primary flying cue.

The RMU docking task is unusual in that it involves the following:

1. The ATS-V is spinning and requires a spinning RMU docking system to be synchronized to null relative motion at the docking interface.

-2. The ATS-V has its longitudinal axis pointing 900 to the sun's rays such that a lighting system is required on the RMU to illuminate the structure (19 inches inside of the ATS aft skirt) to be used for docking.

3. Since the ATS is already in orbit, docking alignment aids can only be installed on the RMU and its ground equipment.

4. Command signals from the ground to the RMU and visual response from the vehicle to the ground take .22.seconds to cover the 40,000 n.m. round trip.

5. Line-of-sight communications for T.V. transmission limit attitude excursions of the RMU in pitch & roll to within + 6 degrees of local vertical.

6. The ATS is spinning about an axis that is .90 off of its geometrical centerline. 1.60 offset was built into the ATS-V model used in the study as a conservative margin.

All of these factors have been designed into the simulator to make the piloting task as real as'possible.

Two-sets of initial conditions (IC) were useAiduring the study. The first set used during the variation of parameters study, is shown in Table I below:

C-1 SD 71-286 Space Division 9D North American Rockwell

Table 1. Initial ConditionVariation of Parameter Runs

Closing Lateral Pitch Yaw Roll X* Y Z Velocity Velocity Attitude Attitude Attitude Range Range Range IC (FPS) (FPS) (Deg.) (Deg.) (Deg.) (Ft.) (Ft.) (Ft.)

1 1.0 0 0 +3 0 50 -10 +10

2 1.0 0 0 -3 0 50 +10 -10

3 1.0 0 +3 0 0 50 +10 +10

4 1.0 0 -3 0 0 50 -10 -10

*Originolly 160 ft. but changed to 50 ft. when established it was unnecessary to start further away.

The second set, used during the final baseline runs (identified by a "B" following the run code number), is shown in Table 2, below.

Table 2. Initial Condition,Baseline Runs

Closing Lateral Pitch Yaw Roll X* Y Z Velocity Velocit Attitude Attitude Attitude Range Range Range IC (FPS) (FPS) (Deg.) (Deg.) (Deg.) (Ft.) (Ft.) (Ft.)

1 1.0 0 +3 +3 0 50 -10 +10

2 1.0 0 -3 -3 0 50 +10 -10

3 1.0 0 +3 -3 0 50 +10 +10

4 1.0 0 -3 +3 0 50 -10 +10

C-2 SD 71-286 d#b space Division North Amercan Rockwell

2.0 SIMULATION EQUIPMENT DESCRIPTION

The simulation equipment consisted of four major parts; the pilot console, the derotation optics system, the television system and the models (ATS and RMU). The initial requirements for each of these systems may be found in Appendix B, Simulation Requirements Document.

2.1 PILOT CONSOLE

The final configuration of the pilot console incorporated two bays, a pilot station and a co-pilot station shown in Figures 1 and 2. The pilot station contained the following equipment items:

1. A 17-inch 525 line Television monitor 2. Rotational hand controller (Apollo block II) 3. Translational hand controller (Apollo block II) 4. Roll attitude indicator + 20 degrees 5. Pitch-yaw attitude indicator 0-20 degrees 6. Latch on indicator light 7. Vertical rate indicator+ 0.1 ft./sec. 8. Horizontal rate indicatr + 0. 1 ft./sec. 9. Range rate indicator-0.5- + 1.0 ft./sec. 10. Minimum impulse switch 11. Rate indicator armed light 12. Rate indicator reset switch 13. RMU light switch

The translation and rotational controller were discrete commands of variable amplitude. The final amplitudes, as established by parameter variation, were .155 ft./sec. 2 and 1.2/sec. respectively. The translation AV was obtained by using the translation controller command signal and a nominal thrust shaping to obtain thrust. Due to the statistical nature (repeatability) of the reaction jet impulse in the minimum impulse mode, a sinusoidal variation of + 10 percent was superimposed on the nominal minimum impulse thrust level. The mintum impulse duration selected was .06 sec.

The co-pilot bay contained­

1. 17-inch 525 line Television monitor 2. Derotation optics switch 3. Despin optics control vernier 4. Latch index display (5" oscilloscope) 5. RMU cage spin up switch 6. RMU cage vernier control C-3 SD 71-286 Space Division North Americn Rockwell

FIGURE 1. RMU PILOT AND CO-PILOT CONSOLE

C-4 SD 71-286 0!

FIGURE 2. RMU CONSOLE DISPLAY LAYOUT

C-5 SD71-286 Space Division North Arnencan Rockwell

The derotation optics and spin cage vernier controls were 2700 potentiometers with +1.5 and +3 RPM command authority respectively. The latch index indication was o]tained by"displaying the differentiated output of two continuous potentiometers on a dual beam oscilloscope. Indexing was accomplished by aligning the transient pulses (one from the ATS model and one from the RMU spin cage) using the ATS as a reference.

2.2 TELEVISION SYSTEMS

Two television systems were used during the simulation. This was necessary due to the prohibitive amount of image smear in the proposed 525 line vidicon system. The system used during the learning and parameter variation runs consisted of an image orthicon RMU camera displayed on two separate monitors, one viewed by the derotation optics camera (525 line) and one viewed by a second 525 line camera. This inter­ mediate step was necessary to maintain the established 525 line system. For the final baseline configuration a lead oxide tube was substituted for the image orthicon, this eliminated the necessity for the intermediate monitor, and made the system highly representative of the spacecraft system. Resolution at the pilot and co-pilot consoles throughout the entire study was 400 lines horizontal and 300 lines vertical.

2.3 DEROTATION OPTICS SYSTEM

The derotation optics system consisted of a 525 line TV camera viewing an 8-inch monitor through a servo driven DOVE prism as shown in Figures 3 and 4. Alignment and sizing of the despun image was obtained by accurate positioning of the 8-inch monitor image. This positioning required both mechanical and electrical adjustments. Align­ ment of the image center of the 8-inch monitor and spin axis of the DOVE prism was established to within a 3/32-inch diameter circle when viewed on the co-pilot monitor. To correct the image inversion due to the DOVE prism, the 8-inch monitor image was electronically reversed.

2.4 MODEL SYSTEMS

The specifications and detailed characteristics of the RMU and ATS models are contained in Appendix B, Simulation Requirements Document. Two modifications were incorporated in the spin cage model. First the cage framework was painted flat black and the latches gray due to their high reflectivity from the RMU camera lights. Second, a shroud was found necessary to shield the spin cage from the sun. Figure 5 is a photo of the spin-controlled docking system and the ATS model.

2.5 CONSOLE RECOMMENDATIONS

The course of the study indicated the console should consist of three bays: a pilot bay, a co-pilot bay, and status engineer bay. As in the study, this console should also be supported by a separate visual system engineer's console (similar to the center console on Figure 6). A list of the equipment items necessary to support both consoles is as follows: C-6 SD 71-286 FIGURE 3. DEROTATION OPTICS SYSTEM

C-7 SD) 71-286 Non Rotating RMU Camera ATS Model

(s Simulator Gimbal System 0 0 Spinning Docking Sse ATransport Rig Derotation Ground Optics Camera Monitor --- t _ .

Ground.------­

0 i / I&'

4" Switch & Vernier .. 3 Docking System Spin 0 01 Switch & Vernier 1 2 3 Co-Pilot Monitor FIGURE 4. VISUAL DISPLAY SYSTEM 0 0 9 Space Division North Ameican Rockwell NOT REPRODUCIBLE

03

FIGURE 5. RMU DOCKING MODEL AND ATS MODEL

C-9 SD 71-286 Space Division North Arnencan Rockwell

FIGURE 6. VISUAL SYSTEMS CONSOLE

C-10 SD 71-286 9Space Division North American Rockwell

RMU CONSOLE EQUIPMENT

Pilot Bay

1. I TV monitor 17-inch 525 line 2. RMU light switch (three position switch, off, single, dual) 3. 1 rotational controller 4. ATS despin initiate switch 5. 1 translational controller 6. 1 minimum impulse switch 7. Roll attitude display + 20 degrees 8. Rate reset switch (resets trans rate displays) 9. Pitch yaw attitude display 0-20 degrees 10. X Trans. rate display - 0.5 to + 1.0 ft/sec 11. Y Trans. rate display - 0. 1 to +0. 1 ft/sec 12. Z Trans. rate display - 0. 1 to + 0.1 ft/sec

Status Engineer Bay

13. Roll gimbal angle display + 50 degrees 14. Pitch gimbal angle display + 50 degrees 15. Yaw gimbal angle display -'50 degrees 16. Roll gimbal angle CMD momentary switch 17. Pitch gimbal angle CMD momentary switch 18. Yaw gimbal angle CMD momentary switch 19. Jet temp monitor (status lights - 12 ea.) 20. Latch status lights (armed, latched, open) 21. Fuel monitor display 22. Latch pressure display 23. Power monitor 24. D.C. Voltmeter with two position switch 25. Mission profile (time) indicator clock

Co-Pilot Bay

26. Event timer (digital shop watch) 27. 1 TV monitor 17-inch 525 line (CONRAC) 28. 1 latch index display (oscilloscope) 29. Derotation optics control

a. Switch (controliswitch) b. Verniercontrol (pot) c. Spin-despin switch (view switch)

C-II SD 71-286 Space Division Nortth American Rockwell

Co-Pilot Bay (continued)

30. Spin cage control

a. Switch (control switch) b. Vernier control (pot)

31. Latch separation switch 32. Latch status lights (armed, latched, open) 33. Latch arm switch 34. TV lens switch (700 FOV 140 FOV)

Visual System Engineer Console

35. 1 TV camera control unit, despin optics, COHU model 3900 (with Dot bar generator) 36. 1 TV camera control unit, RMU camera (dual unit) 37. RMU light status (indicator light for single tube or dual tube operation) 38. Despin optics servo checkout unit 39. D.C. volt meter with ten position switch for systems monitoring 40. 1 17-inch TV monitor 525 line

C-12 SD 71-286 SSpaceNorthAmerican Division Rockwell

3.0 PILOTING PROCEDURES

During the course of the study, piloting procedures were developed from the preliminary procedures defined in Appendix B, Simulation Requirements Document. The final procedures were as follows:

1. Between 200 and 100 ft. of range establish zero DPG gimbal angles (status engineer task).

2. Decelerate terminal rendezvous AV with 5 seconds -X command at 100 to 50 ft. range as indicated by satellite size with respect to stadia on the TV monitor (pilot's task).

3. Gross align with the docking end of the ATS at a stadia range of 10 ft. and null translational rates (2 sec. -X command) to visual zero. Turn on RMU lights (pilot's task).

4. Activate rate meters and minimum impulse translation command switch (pilot's task).

5. Translate to sun lighted edge of the ATS @ .1 FPS and null translation rates with rate meters (pilot's task).

6. Align front and rear edges of the ATS with the monitor crosshair until the front and rear edges appear to oscillate (due to ATS off axis spin) an equal distance either side of the crosshair. Each attitude correction must be followed by a translation back to align crosshairs with the ATS edges (pilot's task).

7. Rotate the RMU docking system such that the pilot can see the ATS center hub while the crosshairs are aligned on the ATS edge (co-pilot's task). The center hub of the ATS is shown in Figure 7.

8. Correct any attitude error in the other axis by observing where the RMU light reflection is with respect to the crosshairs. Command rotation to fly the crosshair to the light reflection such that the crosshair is centered on the reflection. Each attitude correction must be followed by a translation back to align crosshairs with the ATS edge and one ray of the crosshair with the ATS center hub. Verify attitude alignment with backup pitch/yaw meter (pilot's task).

9. Null translation rates and reset rate meters (pilot's task).

C-13 SD 71-286 ' 1 Space Division North Arnencan Rockwell

FIGURE 7. ATS MODEL THERMAL LOUVERS AND CENTER HUB

C-14 SD 71-286 1141 Space Division W North Amencan Rockwell

10. Translate to ATS edqe 900 from first alignment and null translation rates with rate meters (pilot's task). "

11. Check attitude alignment and make corrections as in steps 6, 7 and 8 as required (pilot's task, co-pilot's task).

1.2. Null translation rates and reset rate meters (pilot's task).

13. Translate to center RMU crosshairs on center hub of ATS @ ;t:.I FPS and null translation rates with rate meters. Ask co-pilot for spin-up when ready (pilot's task).

14. Switch on and fine vernier derotation optics to synchronize and derotate TV image on co-pilot's monitor (co-pilot's task).

15. Switch on and fine vernier cage spin motor to synchronize and index the docking system using the index scope to align the RMU roll angle discrete with the ATS roll angle discrete (co-pilot's task).

16. Check despun view for visual index with respect to the ATS Cannon plugs. On verification, ask pilot to complete docking (co-pilot's task).

17. Check.,DPG gimbal angle status and 'inform pilot ifgimbal saturation in roll fs imminent (status engineer's task).

18. Ifnot, correct roll attitude error caused by docking system spin-up (pilot's task).

19. Ifimminent, request roll gimbal desaturation and correct resulting roll attitude to zero (pilot's task).

20. Center crosshairs on ATS center hub. Null translation rates to visual zero. Reset rate meters (pilot's task).

21. Command + X translation to close to docking contact at .15 to .2FPS (pilot's task).

22. Maintain RMU crosshairs on ATS center hub through latching. Check against despun image on co-pilot's console by observing latch con­ centricity and reduced ATS center hub traverse (pilot's task).

23. Latch indicator light "ON"indicates docking complete ifATS image remains fixed.

24. Start despin procedure (co-pilot's task).

C-15 SD 7-1-286 Space Division 11 North American Rockwell

The above procedures were used except for the "statuisengineer" tasks of desaturating the DPG gimbal angles. The procedures can be interrupted at anytime if the pilot decides to repeat the angular alignment process or if, on translation to dock, the lateral miss distance appears to be unsatisfactory:

The sequence of the piloting procedure is important in that attitude alignment is accomplished before docking system spin-up. If the system was spun up first, the gyrodynamics of the spinning mass would cross-couple vehicle attitudes in a direction 900 to rotational commands required during alignment.

The tasks given a co-pilot made a considerable improvement in docking contact accuracy. The task of synchronization and indexing was too distracting to the pilot's primary task of precision flying.

* C-16 SD 71-286 # . Space Division @J North Amencan Rockwell

4.0 STUDY PARAMETERS

4.1 PILOTING PROCEDURES

A baseline piloting procedure was recommended to the pilots for evaluation. Modifications were made during the study without special sets of maneuvers for that pur­ pose. However, 30 learning runs (10 each pilot) provided the primary source of procedure development.

4.2 CONTROL SYSTEMS

Variations of control power in translation consisted of three values of minimum impulse time span, i.e., .03 seconds, .06 seconds, and .12 seconds. Variation of control sensitivity in rotation consisted of three levels of discrete rotational rate command, i.e., .60/sec., 1 .20/sec. and 2.4°/sec. A group of 78 runs were devoted to control systems variations.

4.3 PILOT CONSOLE INSTRUMENTATION AND DISPLAYS

The pilot console configuration and displays were not varied as a parameter in the simulation study but were evaluated and changed to suit pilot comment as much as possible. The primary change from the console described in Appendix Bwas the inclusion of a co-pilot station which provided the derotation optics monitor and docking system spin­ up controls. The pitch and yaw attitude meters were removed and a combined pitch­ yaw meter was installed to be compatible with the RMU/ATS downlink data. The range and range-rate' meters were changed to three Psudo-rate meters.

4.4 VISUAL SYSTEM

Subjective evaluation of visual systems were accomplished during the study. The RMU TV camera used during the major part of the study was an image orthicon which provided a clear non-smear image. During the baseline runs a lead oxide, reduced smear, camera was used and found acceptable.

The sun angle impinging on the ATS was varied from 00 to 220. Results of sub­ jective evaluation were that the 22' sun angle degrades the TV image with too much glare such that proper attitude alignment was difficult. Additional adjustment of the RMU camera failed to improve the image.

A fully open ATS thermal control louvre and a fully closed louvre were evaluated and proven to have negligible affects on the docking problem.

,C-17 SD 71-286 # Space Division North American Rockwell

4.5 BASELINE CONFIGURATION RUNS

After piloting procedures were completed, the control systems parameters were selected, all modifications to the console completed, and the prototype TV systems installed, a series of 30 runs were accomplished to establish statistical data for pilot accuracy as a basis for developing docking system design criteria.

C-18 SD 71-286 AN Space Division Rockwell r4 &Y North American

5.0 SIMULATION MECHANIZATION

The analog computer mechanization for the simulation is shown in Figure 8, sheets 1 and 2. The simulation system was verified by an open loop checkout and a closed loop dynamic verification. The open loop checkout consisted of' verification of the individual subsystem performance paths and was comprised of:

1. Analog static test 2. General purpose visual display interface checkout 3. Television system checkout 4. Spin cage control system checkout 5. Pilot console interface checkout 6. ATS and RMU model verification 7. Tape time delay verification 8. RMU attitude control system verification

The criterion and tolerances for the above open loop checks were established in Appendix B, Simulation Requirements Document.

The closed loop dynamic checkout was composed of two parts, a dynamic verifi­ cation of the analog math model and a man-in-the-loop (qualitative) total system verification. - To support the dynamic analog verification a separate and independent digital model was developed using continuous systems modeling program (CSMP). Given sets of initial conditions were established so as to exercise all portions of the math models. The results of the analog and digital solutions were then compared. In all cases the results were found to be within the readability of the strip charts and CRT printouts. Representative dynamic verification runs are contained in Figures 9 and 10.

The man-in-the-loop verification was conducted by North American test pilots and verified system response to selected commands.

C-19 SD 71-286 DPG SIGNALS 0601-0.I"6 ME. 0000

- ' - . -"i. ....-- . "- ...... - . . . - - -...... -.- ... . - ­ ...... : -- = -_- _. ----- .- - , - ­

D m --- -;.- 5 . 9..

.8

s._,_ _-... . ,__ ,] _ - H. G .4- -V - I I I -

>1 _,l I - __!! K "li _ ' i 12 1 i s I 20

i _ _ 'I I !' ' i

8I 10 12 ['4 16 lS 20 TIME

- _ !_ _ _1

_ _ _ _ _' I " - - - D I -z _

-.s_ _ __ " 'f, I I _ S"1: ! I !I till! ' Si I i I i ; I !I , , -C-2

-.l- e _ ___ _ II I : !i I AS I iS 1 II 4li8S

TI a8S10SD 7128 12 14 Is is 20 9 Space Division NorthAmerican Rockwell

DPG SIGNALS

+1cC______I SD I T ,

AE P4z __

_ __-- ,i l _ _ _ II , , IIi ; iI I I _lLJW! i i I !jWI I lIFI I I I j' J +50 I

I EI i I o -' ',I' ',,' iI h, I t i I[I ii~

I} I; I ,l ! ,! lIlI I.!

itill I~ii il I

S 0fl i __ ' _'I ! t- JH IVi - I'Il

lB I ,1---_---I------1 I - JL0 LJ{ I II I!tI iz !{iz I ! ! I zir - ! I ,,,,. . , I I Il IE,i,' - i

A _E I. I I * 1 a 1111111 II l.....Ii J.i .' ...! . . . 11iI [ [ I I__ -_

0 5 10 15

TiME SECONDS

FIGURE 10. ANALOG TIME HISTORY

C-26 SD 71-286 0 Space Division NorthAmerican Rockwell

6.0 STUDY RESULTS

6.1 SIMULATION RUNS

Approximately 130 dockings were simulated by three pilots in the course of pilot training and "operational" runs. The run log is shown in Table 3. Figures 11 and 12 show lateral miss distances and angular misalignment developed by the pilots in 30 base­ line runs and 15 similar variation of parameters runs. The computed .6845, .9545, and .9974 dispersion envelopes of pilot accuracy are superimposed on the Figures (see next section for a discussion of the statistical treatment).

6.2 STATISTICAL ANALYSIS

The variables included in the statistical analysis and the computer coding are:

1. Y and Z miss distance - RT23Y, RT23Z 2. Rotational fuel -FR 3. Translational fuel - FT 4. Pitch and yaw angles - Theta, PSI 5. Translational velocities - U14T, VI4T, W14T 6. Rotational velocities - P, Q, R 7. Mission time .- T

With the exception of the closing velocity, propellant utilization, and mission time, the distribution of data points for the above variables is approximately normal. The statistical data reduction equations used for the bi-variant variables are as follows:

Arithmetic mean:

A4 i U X, Where N = sample size

xi = value of the t th sample

Absolute mean:

J

C-27 SD 71-286 V Space Division North Amencan Rockwell

Table 3. RMU/ATS Docking Run Log

RUN CODE NO. 01 1 A 11 04 Run No . - I L Dytyof Week Condition No. Pilot Month

Run Variation Code No. RHC THC Display ProcedureRemarks 011A1104 .3°/sec No mini All Both dock and Good docking pulse index 022A1 104 . " " Abort, skirt contact 033A1 104 J, " ". Large miss 044A1 104 " meter OK 051A1104 High closing vel. 062A1 104 Abort, skirt contact 073A1104 OK 084A1104 1.2/sec .03 sec All Split task OK 091AI104 OK

END TRAINING RUNS - BEGIN VARIATION OF PARAMETERS RUNS 113AlII18 .60/sec .03sec All Splittask OK 124A1118 OK 131Al118 OK 142A1 118 OK 153A1118 OK 164A1118 2.4e/sec .03 sec OK Att. drift 171A1118 OK 182A1118 OK 193A1 118 OK 204A1118 OK 211A1118 1.2°/sec .3 sec OK 222A]118 " " " " OK 233A 1119 O K 244A1 119 Bad run X too high 251A1119 OK 262A1119 1.20/sec .06 sec OK 273A1119 O K 284A1119 OK 291A1119 OK 303A1119 " OK 314A1120 1.2/sec .12 sec All Split Task OK 321A1120 .06 sec None Split Task OK 332A1120 ". OK END VARIATION OF PARAMETERS RUNS - BEGIN BASELINE RUNS - FLY TO LIGHT WEDGE

NOTE: RHC = Rotational Hand Controller Rate Command Discrete THC = Translational Hand Controller Minimum Impulse Time Span

C-28 SD 71-286 SpaceDivision NorthAmencan Rockwell

Table 3. RMU/ATS Docking Run Log (Continued)

Run Variation Code No. RHC THC Display Procedure Remarks 011A1201B 1.20/sec .06 sec All Split Task OK 022A1201B OK 033A1201 B " " I OK Cage process required 044A1201B OK 051A1201 B OK 062A1201 B OK 073A1201 B OK 084A 1201 B OK 091A1201 B OK 102A 1201BBt OK END PILOT A RUNS, BEGIN TRAINING RUNS PILOT B 011I1105 .60/sec No mini All Dual Task OK pulse 022B1 105 Large miss 033B1105 Large miss 044B1105 " " only OK 051B1105 1" OK 06281113 1.20/sec .03 sec All Split Task OK 073B1113 OK 08481113 OK 091B1113 OK 102B] 116 OK

END TRAINING RUNS - BEGIN VARIATION OF PARAMETERS RUNS 113B1116 .60/sec .03 sec All Split Task OK 124B1116 OK 131B1116 OK 142B1116 OK 153B1116 ' OK 1641116 2.40/sec OK 171 B1116 OK'

182B1116 It OK 19311116 OK 204B1116 I Off in Y' 211B1116 1.20/sec .06 sec O K 222BI116 OK 233B1117 OK 24411117 OK 251BI 117 OK 26281117 1.20/sec .12 sec OK, No vernier on translation 273B11J7 .06 sec None OK 28481117 OK 291B1117 OK END VARIATION OF PARAMETERS RUNS - BEGIN BASELINE RUNS - FLY TO LIGHT WEDGE

C-29 SD 71-286 0 Space Division North Amencan Rockwell

Table 3. RMU/ATS Docking Run Log (Continued)

Run Variation Remarks Code No. RHC THC Display Procedure

01181130B 1.20/sec .06sec All Split Task OK 022B1130B OK 03331130B OK 044BI1308 OK 051 BI 130B I O K 062811308 OK 07381130B " First of morning 084B1202B OK 09181202B OK 102B1202B "OK 11181202B 22°/sec Sun, Open Louvers, Goodrunbut Iight wedge n/g 12181202B Same but w/RMU lights off RMU lights have no effect 131812028 00Sun, Louvers Closed No effect on baseline END PILOT B RUNS - BEGIN TRAINING RUNS PILOT C 011C109 .60/sec .03sec All Dual Task OK

022C1104 "t " I 2-inch miss 033C1109 "t OK 044C1109 " OK 051C1109 3-inch miss,33 miss 062C1109 "" 20 miss 073C109 " " .30 miss 084C1109 "I....N/G * too high 091C1109 OK 102C1109 .. N/G X too high 113C1111 Split Task OK 124(1111 "t OK

END TRAINING RUNS - BEGIN VARIATION OF PARAMETERS RUNS 131C111 1.20/sec .03sec All Split Task OK 142CI1111 " " I9,Y = 30 153C1111 9 = 2.50 164C1111 " " OK0I 171C1111 OK 182C1111 2.40/sec 0,P = 2.60, 1.80 193C1111 OK 204C1111 OK 211C1112 OK 222C1112 It ". OK 233C112 1.20/sec .06 sec t " End run before visually docked 244C1112 " " " " OK 251C11 T2 " J " I OK, loss of index simulated 262C1112 " " " t OK 273C1112 " " " " OK, loss of index simulated 28401112 . 12 sec " " OK but too sensitive 291C1112 .03 sec None " OK 302C1112 " " " " OK 313C1112 I " " OK

C-30 SD 71-286 0 SpaceDivision North Amencan Rockwell

Table 3. RMU/ATS Docking Run Log (Continued)

Run Variation

Code No. RHC THC Display Procedure Remarks

END VARIATION OF PARAMETERS RUNS - BEGIN BASELINE RUNS - FLY TO LIGHT WEDGE

01 IC1124B 1.20/sec .06 sec All Split Task OK 022C1124B 0 K 033C1124B O K 044C1124B " I " T Yaw DPG's saturated, OK 1" 051C1124B "1 IT . 062C1201B T T OK, Uses 5 sec -X at 50' 073C1201B " I OK 084C 1201B OK 091C1201B " OK0I 102C1202B " " OK

C-31 SD 71-286 j ~Space Division North American Rockwell

-Z .6875 "random" error "Total" error probability circle -2 probability circles

.9974 .9545

- .6877

\j\

-2 "Y 0 2 + -y -21 'o Joo@ , 6+

; 00t /

• i/ ' ­

+2­

* Annlicable variation of Points do not include parameter data points 0.46 inches transverse (15 runs) +Z error caused by ATS-V 1.6 degree wobble o Baseline data points (30 runs) The arithmetic mean and the probability circles Arithmetic mean are based only on base­ line data points (30 runs)

Figure 11 Pilot Lateral Miss Distance (Inches) Accuracy

C-32 SD71-286 #11 Space Division North American Rockwell

.6875 "random" error ­ "Total" error probability circle probability circles -2 .9974

.9545

._~ .. 6877

00 '0,

-2 +-1+

-

+1 /

+2-

* Applicable variation of Points do not include parameter data points 1.6 degree ATS-V wobble (15 runs) misalignment

G Baseline data points The arithmetic mean and (30 runs) the probability circles are based only on base­ Arithmetic mean line data points (30 runs)

Figure 12 Pilot Angular Misalignment (Degrees) Accuracy C-33 SD7- -286 @ Space Division North American Rockwell

Sample set standard deviation: A r -= XL i N-I

Four bi-variant variables were considered. They consisted of translational miss distance, radial velocity, angular miss distance, and angular rates . Assuming the 6i-variants unbiased and uncorrelated (as a first approximation), the probability density function (pdf) for each of the four variables is given by a normal bi-variate error distribution of the form:

A circle of radius (d) centered at the point (X, Z) enclosing (P)percent of the data is obtained by using the following equation (reference 3):

where X and Z are the arithmetic means of the Xand Z components and P is the desired percentage. Three values of Pwere used in the calculations, .6827, .9545, and .9974 yielding multipliers (K) of .74, 1.24, and 1.73, respectively. Since the center of the circle described above is offset from (0, 0) by the total radius is obtained by:

as shown in Figure 13. This, in effect, provides a conservative estimate of total miss distance including both bias and random variation. For illustration, the resulting probability envelopes of pilot accuracy for lateral and angular misalignments are plotted on Figures 11 and 12, in the preceding subparagraph.

A data reduction program was developed to convert the analog outputs at the end of the simulation runs to engineering units and to calculate the statistics for the simulation. The results for baseline runs are shown in Tables 4 through 8. The printout code for the variables listed is given in Appendix B, Simulation Requirements Document. Axial velocity values were highly skewed and, therefore, the value of the axial velocity listed in the lower portion of the printout is the maximum value seen over the data set and not the mean plus 3 sigma. The two sigma and 3 sigma values referenced in the printout are for probabilities of occurrence of .9545 and .9974, respectively, rather than 2 or 3 sigma values in the strict sense of those terms.

C-34 SD 71-286 Space Division OD NorthAmerican Rockwell

: d =-K(or. ,,+or) " +

Figure 13. Miss Distance Probability Envelope

The printout in Tables 4 through 8 for radial velocity and angular rates presents the data in accordance with the above equation for total radius. For lateral miss dis­ tance and angular misalignment, a factor was added to the equation to account for ATS-V wobble induced offsets relative to the latches. Therefore, the equation for these variables are:

Lateral miss - t, t % -- c

Angular miss = -t K 6e +)4--.6

The lateral miss and angular miss data presented, then, exceed pilot accuracy probabilities by the quantities shown and provide latch misalignment tolerance design criteria.

'C-35 SD 71-286 TABLE 4. COMBINED TRAINING RUNS - 30 SAMPLES

XBAR XABS STANDARD DZVIATION.MLAN--. S-IG3M RT23Y #75440E 00 .19136E Ci ,37839E O1 ,12106E'02 RT23Z -.71320E o0 17356E Ci *24969E 01 .677hSE 01. FT *46215E 00 .46215E CO ,23052E 00 i1537E 01 FR '21952E-o *21985E-Ci o2307oE-01 *91161E-O1 T *53108E 03 ,531c8E C3 *19696E 02 "59017E 63 THETA 18462E 01 *18558E Ci 15346E 01 ,64499E 01_ PSI *77573E 00 .81859E CO 951685E 00 *23263E O1 Ui4T ,24 85oE 0i ,26250E C1 ,25579E 0i o10159E o2 V14T .12152E 00 .16984E CO *29813E OO -10159E 01 WI4T -*73600E-02 033344E CO *57024E 00 .17034E 01 p -.12988E-02 #33234E-c2 960471E-02 -16843E-01 O ,15089E-02 ,15c89E-c2 910692E-02 .47164E-02 R *17763E-02 *17763E-c2 .89468E-03 *44603E-02 C-) MEAN + 3SIGMA MEAN + 2SIGMA LATERAL MISS *12333E# C2 .93052E 1 Inches FUEL TOTAL *90C12E CO .77789E OO Lbs.

TIME - *59017E C3 000E 03 Sec-.-----.

ANGULAR MISS - .71412E Ci ,61524E 01 Deg. RADIAL VELeCITY *16197E Cl *12011e 01 Inch~ec. z W AXIAL VEeCITY .13842E C2 Inch/Se 0 > m LATERAL ANGULAR RATES *57183E-c2 *47717E-02 Deg/Sec. M.<

*STOP* 46266024 _ _. 0

IMPROPER CONTROL CARD SEQUENCE &EeF TABLE 5. BASELINE RUNS PILOT "A" - 10 SAMPLES XBAR XABS RT23Y STANDARD DEVIATION MEAN + 3SIGMA RT23z ,26160E,96000E-02 00 .40080E,32880E CO ,39391E937976E CO..o0 914433E,UAS8SE.0I 01 FT .62623E 00 ..q6262 ....---_2.U3L C.. .nSA2. onCQ FR *16340E-01 *16340E-d1 911614E-02 019824Ew"0. T *51961E 03 .51961E 03 ,11401E 01 ,52303E..03... THETA *18370E 00 ,18370E Co o77763E-o1 941699E 00 PSI *15712E 00 .25544E o0 .35993E 00 . .2369E Ol - U14T *16512E 01 ,16512E C! *53992E 00 ,32710E 01 V14T '80400E'02 t55080E-C1 ..... -22226E...cL W14T -13440E-01 977280E-C1 10195E 00 *29242E 00 P ,12033Em02 ,48705E-c2 ,69207E-02 ... 21965EO.l Q .,68760Eo03 ,68760E"C3 #84559E-03 #18492E-02 R .,2292oE-03 o57300E-C3 977349E-03. 20913E-02

LATERAL MISS * ,20563E 0l *16834E 01.... FUEL TOTAL .#79128E 00 o74523E OQ. Lbs.....

TIME a o5a03E 0...... v52.1& .2 c.

A ANGULAR MISS = 25968E l o23858E 01 Deg. RADIAL VELOCITY ; .32045E 00 .923529K 0.0 Inch/Sec. c AXIAL VELOCITY .#2634QE ...... lcks .. zZC/z 04 0o . . LATERAL ANGULAR RATES 35177E-Q ?. . 7a3 E .2 aA .. . -

*STOP* 77774123 0l<

0 IMPROPER CONTR8L. CARD SEQUENCE ...... --...... AEOF TABLE 6 BASELINE RUNS PILOT "B"-10 SAMPLES XABS STANDARD DEVIATION MEAN + 3SIGMA RT23Y -#37440EXBAR 00 ,37920E CO ,34072E 00 *64776E 00 RT23z ,28080E O0 ,53280E co #57684E 00 .20i13E 01

FT .46384E Oc t46384E.-c0. . ±5iiLZEtc$ . . .3353. .0. FR 916520E-01 ,16520E01. *49086E-02 931246E-01 T #59156E 03 .59156E 03 912699E 03 ,97253E.03. THETA ,10085E-01 ,35847E CO ,41894E 00 -12669E 01 PSI *53186E 00 *65219E CO #66240E 00 t25191E 01­ U14T ,15j74E 01 *15174E ci *96350E 00 *44079E O VX4T ,4l880E-01 *84360E-ci... 11411E.-Q .....,c3842OE.O.. W14T -#27600E-01 *39360E-Ci ,46582E-01 *11215E 00 P -,29223E-02 ,41829E-c2 955321E-02 ,13674E-0.1 a ,OOOOoE 00 *OOOOOE CO 9O0OOOE 00 *O00OOE 00 R *11460E-02 *11460E-C2 #81034E-03 *35770E-02

MEAN + 3SIGMA MEAN + 2SIG A. .. LATERAL MISS ,25108E9 Cl ,20685E 01 __ Iacha___

FUEL TOTAL v *57389E CO ,54323E 0o Lbs.....

TIME Rp)7a53E 03 t8.t55.L ... Se . a, ANGULAR MISS 9,39973E Cl .34761E O0 . De.

RADIAL VELOCITY * *32734E CO 924989E 00 inch/Seq. 0 AXIAL VELOCITY .26400E C1 . . Z (

LATERAL ANGULAR RATES x ,25438E.C2 ,2 53q E~t- . e./S c._ 9>

3 *STOP* 77774123 0 J 0 IMPROPER CONTROL CARD SEQUENCE AEOF TABLE 7. BASELINE RUNS PILOT 'C"- 10 SAMPLES XBAR XABS STANDARD DEVIATION MEAN + 3SIGMA RT?3Y .,91200E-01 934080E 00 943628E 0..­ .. 2176E0± RT23z #49080E 00 .52200E CO o47263E 00 -19087E 01 FT4?o34E 00 42,03a4E.co...u.E.n ~ ~ 52-0 FR *13035E-01 *13035E-01 *34029E-02 ,23244E-01 T #53805E 03 ,53805E 03 *85852E.0-. .95E._03... THETA .26690E O0 o28249E 00 .19989E 00 #86657E 00 PSI #14096E 00 .46413E 00 , 6,5403E2 .0... _t 2103QE01-_ U14T 990840E 00 *90840E 00 .25954E O0 ,16870E 01

V14T ,22200E-01 . 51480ETOL ..... ,_61aZE0 2o7sE,0­ WX4T t58800E-02 ,60i20E-0I @96878E-01 *29652E 00

P #45840E-03 ,43548E-c2 - .57032E!02-. .At_756aE .01 Q o57300E-04 ,57300E-04 o18120E-03 ,60090E-03 R -.40110E,03 *10887E-0? o17516E0--..--- -8537E-02_.

MEAN + 3SIGMA MEAN + 2SIGMA LATERA MISS v o25271E 01 ,.2089QE.01 Inches

FUEL TeTAL i .6191$E CO 956363E.0 ..--.-- Lbs...

TIME . 7 .79561E 0. .. . .ZQZ5E_ ... .. Sec. ANGULAR MISS 9 *33748E Ol .29633E 01 . Deg.

M RADIAL VELOCITY .29690E 00------22036E_O0 ...... L[oh/5e._. coAXIAL VELOCITY .±OW.QCi- za

LATERAL ANGULAR RATES ; .37392E-02 *28076E-Q ..Peg/Sec, -D > fl< <- *STOP* 77774123 Cn

IMPROPER CONTROL CARD SEQUENCE 4EeFm TABLE 8. COMBINED BASELINE RUNS-30 SAMPLES

XBAR XABS SiANDAiD.PEVIAT.I.EAN.-± 3 .1A RT23Y -915200E 00 *34960E CO *40889E 00 *10747E 01 RT23z *34440E 00 *48520E CO *48155E O0 *17891E 01 FT #50347E 00 *50347E 00 ,12582E 00 *88092E 00 FR ,152 9 8E-01 o15298E-01 *3761oE-02 *26581E-0 T 954974E 03 .54974E C3 *90862E 02 ,82232E 03 THETA ,15356E 00 927489E 00 *8388E . . * QQ52E,0O PSI 9276642 00 *5725E 00 ,585552 00 ,20333E 01 U14T ,13590E 01 -135oE C1 ,71245E 00 #34964E 01 V14T 924040E-01 63640E-Cl *84638E-01 .27795E 00 W14T -.11720E-01 #58920E-01 #83706E-01 ,23940E 00 P -.,42020E03 .44694E-c2 #61474E-02 #18022E-01 Q q2101oE-03 *24830E-03 *59211E.03... . 15662E-02 R *17190E-03 *93590E-c3 °13555E-02 .42385E-02

MEAN + 3SIGMA MEAN + 2SIGMA LATERAL MISS ,23725E Cl ,19433E 01 Inches

FUEL TOTAL 972722E CO ,66477E 00 - £s.

TIME *82232E C3 .73146E 03 Sec. ANGULAR MISS * @34162E Cl ,29971E 01 Deg. RADIAL VELOCITY ,31714E9 CO ,23600 "00.'- Inches/Sec. 0' AXIAL VELOCITY .26400E C1 Inches/Sec. n 0) = 0 LATERAL ANGULAR RATES *36312E-C2 *26924E-02 Deg/Sec.' P

*STOP* 77774123 . 0 0 IMPROPER CONTROL CARD SEQUENCE AEBF Space Division 0D NorthAmencan Rockwell

7.0 DISCUSSION OF RESULTS

7.1 EFFECT OF LEARNING

Since the development of piloting procedures continued throughout the study, only a relative measure of their effects can be demonstrated numerically. Table 9 presents statistical docking conditions generated by 30 learning runs compared to the final set of 30 baseline runs using fully developed procedures and RMU systems. Note that propellant utilization increases with flying precision. No other attempt was made to establish the numerical effect on contact condition due to changes in procedures, training, or pilot fatigue.

7.2 EFFECTS OF ROTATIONAL CONTROL SENSITIVITY

The output of the pilot's rotational hand controller is a constant level signal sent to the RMU as a discrete. The magnitude of the discrete represents an angular rate command to the RMU. The pilot's ability to use the RMU rotational control system with precision depends on the magnitude of this rate command. The pilot refers to this as control sensitivity. In reality, sensitivity includes not only the magnitude of the command but the vehicle response to the command and how much controller deflection was required at what controller force. To make the most efficient use of the short schedule of the simulation, it was assumed that the Apollo block II controller, shown on Figure 14, was aaceptable as to command deflection and deflection force. It was also assumed that the dual purpose gyro control response to command was acceptable. Thus, a variation of command magnitude was accomplished assuming all other components of sensitivity were known acceptable. Rotational rate commands of .60/ sec. felt too slow to the pilots. Rate commands of 2.4°/sec. provided too much attitude overshoot. Rate commands of 1.2°/sec. were agreed upon by the pilots and verified by the numerical data for rotational hand controller (RHC) variation shown on Table 10 in terms of improved attitude alignment (see blocked numbers).

7.3 EFFECTS OF TRANSLATIONAL MINIMUM IMPULSE

During the early training runs, a conventional translational acceleration command system was mechanized which provided .193 ft./sec. 2 for as long as the translation con­ troller was deflected. The tendency to overcontrol caused a reduction to .155 ft./sec.2 which amounts to 1.2 lb. thrust per RMU jet. The reduced control power improved control of translational velocities during gross maneuvers, but to fine tune the translational rates to a visual null, a much smaller control power was needed. Since 1.2 lb. jets were a lower limit for space-qualified reaction jets, a minimum impulse mechanization was2 developed. The minimum impulse limits each controller deflection to a .155 ft./sec. command pulse of shorter duration than human response could generate. The task was

C-41 SD 71-286 TABLE 9. TRAINING RUNS VS. FINAL RUNS

MISS DISTANCE iLATERAL VEL. CLOSING VEL. ANGULAR MISS ANGULAR RATE PROPELLANT (INCHES) (INCHES/SEC) (INCHES/SEC) (DEGREES) (DEGREE/SEC) (LBS) NUMBER VARIATION MEAN LIMIT MEAN LIMIT MEAN MAX MEAN LIMIT MEAN LIMIT MEAN MAX SAMPLES

Training 1.91 9.3 .333 .20 2.62 13.8 1.85 6.15 .0018 .0048 .484 .746 30

Final .485 1.94 .0636 .236 1.35 2.64 .457 2.99 .00094 .0027 .519 1.45 30

0

Mean = Absolute Value Average of Worst of Two Axis.

Limit = Envelope of pilot accuracy encompassing 95.45% of all docking attempts. Miss distance limit includes an addi­ tional 0.46 inches to provide for ATS-V wobble induced 0offset; angular miss limit includes an additional 1.6' to provide for ATS-V wobble induced misalignment. a, Max = Highest value seen of a statistically skewed parameter. S

0 *0 >0

D

0

0

SD 71-286 ' Space Division North Amecan Rockwell

NOT REPRODUCIBLE

Apollo Block II Rotation Controller

Apollo Block II Translational Hand Controller

FIGURE 14 HAND CONTROLLERS

C-43 SD 71-286 Table 10. RMU Docking Control System Variations

MISS DISTANCE LATERAL VEL. CLSING VEL. ANGULAR MISS ANGULAR FATE PROPELLANT NMsn (NcHES) (IocHs/sEc) (ImCES/SEC) (E (DEGE /SEC) (LBS SAWLES VARIATION MEAN LIMIT MEAN LIMIT MEAN MX MEAN LIMIT MEAN LIMIT MEAN MAX 0 RHO = .6 /Sec .567 2.21 .092 .297 1.90 3.52 .932 4.34 .0024 .0060 .539 1.48 15

RHO = 1.2°/Sec .585 2.44 .075 .295 1.72 3.07 .783 r.24 .0025 .0080 .62 1.51 15

0 RHO = 2.4 /Sec .557 2.35 .0706 .227 1.86 3.63 1.47 6.24 .004 .0139 .534 1.49 15

Tile - .03 See .585 2.44 .075 .295 1.72 3.07 .783 4.24 .0025 .0080 .62 1.51 15

THe = .06 Sec .510 F28 .066 .198 1.84 3.80 .749 3 .O018 .0037 .528 1.48 15

THC = .12 Sec .640 2.60 .108 .296 1.59 2.68 .698 3.84 .0021 .00306 .564 1.46 3

For all RHC variations THC = .03 seconds Mean =Absolute Value Average of Worst of Two Axis.

For all THC variations RHC = 1.2 /sec. Limit = Envelope of pilot accuracy encompassing 95.45% of all docking attempts. Miss distance limit includes an additional Note: THC =Translational hand controller minimum pulse duration 0.46 inches to provide for ATS-V wobble induced offset angular miss limit includes an additional 1.6 to provide for RHC Rotational hand controller rate command ATS-V wobble induced misalignment.

Max Highest value seen of a statistically skewed parameter.

0 0 0$ Space Division North Amencan Rockwell

Then to establish the pulse width that would provide the most precision at RMU docking. Three levels of minimum pulse duration; .03 seconds, .06 seconds and .12 seconds, were investigated. Mini-pulse durations of .03 seconds were favored by two pilots and .06 seconds by the remaining pilot.

The numerical data shown for translational hand controller (THC) variations on Table 10 favors the .06 second minimum impulse. Accuracies in both miss distance and angular alignment were improved (See blocked numbers).

The .J2second mini-pulse approaches human response capability and was deter­ mined unacceptable after a single run per pilot.

7.4 RMU BASELINE RESULTS

After determining the translational and rotational control system requirements, the simulation was modified to incorporate a spacecraft prototype TV camera consisting of a low smear lead-oxide tube. Using the developed piloting procedures and console con­ figuration, 30 runs were devoted to collecting statistical data for docking system design criteria. Each pilot was to fly 10 runs. The statistical data shown in Table 11 indicates a continued improvement over previous data. Comparison of pilot performance shows pilot "A" to be slightly more accurate in miss distance than the other two. This may be an indication that pilot "A" was using the mini-pulse of .06 seconds that he liked, whereas the other two pilots would have preferred .03 seconds mini-pulse.

A good compromise would be to place a tolerance band on the spacecraft mini­ pulse insuring .04 to .06 seconds, or recommend two levels that can be selected by individual preference.

7.5 EFFECTS OF CONSOLE INSTRUMENTATION

A small number of runs were attempted with cockput instruments covered to de­ termine if precision is degraded. These runs were accomplished after the baseline runs at the highest possible level of training.

Table 12 indicates a degradation of performance even considering that the sample size limits the statistical validity. Of particular interest is the degradation of attitude alignment by approximately 1.5 degrees. This could indicate dependency on the pitch­ yaw meter by one of the pilots and is further verified by examination of individual run data. This being the case, it is recommended that accuracy of the pitch-yaw meter information be at least within + 1.0 degrees. Further training may eliminate this dependency.

.C-45 SD 71-286 TABLE II. RMU BASELINE DOCKING RUNS

MISS DISTANCE LATERAL VEL. CLOSING VEL. ANGULAR MISS ANGULAR RATE PROPELLANT NUMBER (INCHES) (INCHES/SEC) (INCHES/SEC) (DEGREES) (DEGREE/SEC) (LBS) SAMPLES VARIATION MEAN LIMIT MEAN LIMIT MEAN MAX MEANI LIMIT MEAN LIMIT MEAN MAX

Pilot A .401 1.68 .077 .235 1.65 2.63 .255 2.39' .00069 .0027 .78 1.45 10

Pilot B .532 2.07 .0844 .250 1.517 2.64 .652 3.48 .00114 .0022 .48 1.08 10

Pilot C .522 2.09 .0601 .220 .908 1.31 .464 2.96 .0011 .0028 .43 1.35 10

Total Baseline .485 1.94 .0636 .236 1.35 2.64 .457 2.99 .00094 .0027 .52 1.45 30

a,

Mean = Absolute Value Average of Worst of Two Axis.

Limit = Envelope of pilot accuracy encompassing 95.45% of all docking attempts. Miss distance limit includes an addi­ Ntional 0.46 inches to provide for ATS-V wobble induced offset; angular miss limit includes an additional 1.60 to provide for ATS-V wobble induced misalignment. Max = Highest value seen of a statistically skewed parameter. z cn o CD

n.< 0

0

SD 71-286, TABLE 12. EFFECTS OF COVERED INSTRUMENTS

MISS DISTANCE LATERAL VEL. CLOSING VEL. A NGULAR MISS ANGULAR RATE PROPELLANT NUMBER (INCHES) (INCHES/SEC) (INCHES/SEC) (DEGREES) (DEGREE/SEC) (LBS) SAMPLES VARIATION MEAN LIMIT MEAN LIMIT MEAN MAX MEAN LIMIT MEAN LIMIT MEAN MAX

Blind .483 2.28 .06 .2 1.62 3.46 .93 4.46 .0023 .004 .52 1.45 9

Baseline .485 1.94 .0636 .236 1.35 2.64 .457 2.99 .00094 .0027 .52 .45 30

4 Mean = Absolute Value Average of Worst of Two Axis.

Limit = Envelope of pilot accuracy encompassing 95.45% of all docking attempts. Miss distance limit includes an addi­ tional 0.46 inches to provide for ATS-V wobble induced offset; angular miss limit includes an additional 1.6' to provide for ATS-Vwobble induced misalignment. (A Max = Highest value seen of a statistically skewed parameter. N0Z S z U o 0M 3m

0 Space Division North American Rockwell

8.0 PILOT DATA

8.1 RMU RESEARCH PILOT BACKGROUND

J. W. COX (Pilot "A")

Mr. Cox has twenty years experience as a Naval aviator and test pilot. He has flown a wide variety of aircraft from high performance lets to propeller driven transports and has approximately 5000 hours of flight experience. He has participated in flight simulation in the Apollo Mission and Hardware evaluators, the Lunar Flying Vehicle, Space Shuttle and Remote Maneuvering Unit. Mr. Cox is a graduate of the Navy Test Pilot school.

R. 0. RAHN (Pilot "B")

Mr. Rahn has seventeen years experience as an experimental test pilot and has been a pilot for 31 years. He has flown over 60 different types of aircraft and has approximately 6500 hours of flight experience. He has participated in Flight simulations in the Apollo Mission and Hardware evaluators, the Lunar Flying Vehicle, Space Shuttle and Remote Maneuvering Unit, Mr. Rchn is a graduate of the Air Force Test Pilot School.

P. 0. HARWELL (Pilot "C")

Mr. Harwell has twenty-three years experience as a Naval aviator with approximately 6500 flight hours in some 70 different airplane models. He is a graduate of the U.S. Naval Test Pilot School, served two tours as a test pilot, and holds a Commercial license with single, multi-engine, instrument, and helicopter ratings. He has six years experience as an Apollo Research Pilot and has participated in numerous engineering simu­ lations including: Apollo entry, docking, failure'effects; Skylab failure effects; Space Station docking; Advanced Programs Lunar Lander, two previous RMU studies ...... etc.

8.2 PILOT COMMENT

The following was extracted from the pilot's report, IL 642-70-2162, and contains a description of a recommended console instrumentation layout:

CC-49 SD 71-286 Space Division NorthAmerican Rockwell

4.0 RESULTS AND DISCUSSION

4.1 The original configuration evaluated included a one-man operator concept, with control moment gyros (CMG) for attitude control, one pound RCS thrusters capable of compensating for CMG gimbal saturation, THC output of 0.155/sec 2 and RHC discrete rate output of 0.30/sec.

4.2 It was immediately apparent that a minimum impulse (MI) mode for Y and Z translation was required to provide the control refinement necessary to accomplish the docking task. A modification was made to provide the pilot with a selectable 30 millisecond (M/S) minimum impulse mode for Y and Z axes, as well as the acceleration mode mentioned above. It should be noted that X translation control was always in the acceleration mode.

4.3 During the development and training phase, THC minimum impulse widths of 30, 60, and 120 MS as well as RHC output discrete rates of 0.3, 0.6, 1.2, and 2.4 0/sec were evaluated. Complications associated with the one-man operator concept, such as target fixation, undetected target drift while observing the despun dis­ play, time-critical switch selections and a general feeling of hurried task accomplishment lead to the development of a two-man console. Sharing time-critical workloads resulted in a significant improvement in the overall task accomplishment and a marked reduction in docking errors.

4.4 Prior to the production run series, it was determined that the baseline control authority for the simulated RMU would include two THC modes; acceleration command at 0.155'/sec 2 and MI of 60 MS as well as RHC output discrete rate of 1.2 0/sec. Future studies might prove that something less than 60 MS will be de­ sirable for the MI width.

4.5 Four different initial conditions (IC) were used for the majority of the runs. On about five runs, the IC separation distance was moved out to 160 ft. It was determined that, under the conditions of the study, target acquisition at the greater distance was feasible, however, the time required to complete each run was significantly increased. Since the increased time was not considered to be productive, the 50 ft. IC was used for the remaining runs.

C-50 SD 71-286 Space Division 0D North Amencan Rockwell

4.6 The scope of the study did not include a failure effects analysis. However, on a few runs during the early phase, the range meter and/or 11 meters were covered during the entire run. As pilot experience and proficiency increased, other subtle cues were noted in the visual displays and greater flexibility was realized from procedural changes in the manner of utilizing the available meters. A wedge-shaped light pattern on the docking interface of the ATS-V, created-by reflected rays from the RMU lights, was noted to furnish an additional check of pitch and yaw attitude error depending on whether the pilot was sighting down the "12-O'clock or 9-O'clock" edge of the target. Cross­ checking the "Co-pilots" despun display during the last 10 ft. of closure presented the pilot with an additional indication of Y and Z alignment and/or pitch and yaw attitude errors. Utilizing the horizontal and vertical rate meters as "pulse counters" when operating in the Ml mode made starting and stopping translation maneuvers significantly less difficult. A more detailed explanation of the utilization of these additional cues is covered in Paragraph 4.7.

4.7 The procedures and techniques discussed below will refer to those employed during the production run phase with an all-up RMU being flown by a two-man crew. When the run commenced, Y and Z rates were initiated with the THC in the acceleration mode so as to start the target to drift toward the zero translation error reference mark ("cross-hair") on the pilot's CRT display. Immediately there­ after , the closure rate was reduced by a three to five second -X command. The target drift was stopped when translation errors were near zero and closure continued until the distance between the two vehilcles was approximately 10 ft. as evidenced by scribe marks on the pilot's CRT which spanned the ATS-V outside perimeter. At this point, all rates were nulled as accurately as was possible in the acceleration mode. Up to this point in the problem all pilot information had been gleaned from the CRT as the display meters were not activated until all rates were nulled. When the meters were selected ON, translation output signals were integrated to drive the vertical, horizontal, and range rate displays. The MI mode was selected and a -Y translation was commenced while using the horizontal rate meter as a "pulse counter" to aid the pilot in establishing an acceptable rate as well as stopping the translation maneuver. During the maneuver, the vertical rate meter was used to maintain the geometric center of the ATS-V on the horizontal reference line on the pilot's CRT. When the 9-O'clock position on the target was at the center of the pilot's CRT,' the rates were again nulled while exercising extreme care to keep the geometric center of.the ATS-V on the horizontal reference line that passes

S7C-12 SD 71-286 Adb Space Division @T North American Rockwell

through the center of the CRT. The pilot would sight down the 9-O'clock edge of the ATS-V and correct any relative yaw error between the two vehicles. The wedge of reflected light, mentioned in Paragraph 4.6 above, was noted to evaluate any existing relative pitch error between the two vehicles. Using a fly-to convention, the horizontal reference line on the pilot's CRT was superimposed through the center of the wedge of light as shown in Figure 15. The pitch and yaw attitude error meter was checked and with very minor corrections at this point, the meter could usually be zeroed. However, to further verify the precise­ ness with which total alignment had been accomplished and to eliminate the possibility of an erroneous meter indication, a +Y and -Z translation was initiated. The vertical and horizontal rate meters were again used as "pulse counters" to start and stop the maneuver. When the 12-O'clock position on the target was at the center of the pilot's CRT, pitch error was checked by ensuring that the vertical reference line through the center of the CRT was superimposed through the center of the wedge of reflected light. A +Z translation maneuver was commenced to Fly the "cross­ hairs" on the pilot's CRT to the center of the ATS-V docking inter­ face. With all rates nulled and the distance between the two vehicles about 10 ft., the meters were selected OFF and back ON to reset them to zero. During the entire run the co-pilot was using his vernier control to rotate the docking cage left or right sufficiently tb prevent the latches from masking key reference points on the target vehicle. At this time the co-pilot would select the despin optics switch to ON and use his vernier control to null any apparent rotational motion of the ATS-V. He would then spin-up the docking cage and use his vernier control to index the docking latches with respect to cannon plugs located near the docking interface on the target. A separate oscilloscope was provided to display the output signals from a sun sensor on the ATS-V and a magnetic pickup on the rotating docking cage. When these two signals (blips) were coincident, the relative orientation of the ATS-V and the docking latches was correctly indexed and the co-pilot would give the pilot a "GO" to dock, The pilot vould initiate a +X translation to establish a closure rate of about 0.1 ft/sec. Using the MI mode, Y and Z translation rates and errors were maintained at effectively zero. In addition to the information available on the pilot's CRT, the center of the co-pilot's despun view could be cross-checked to get additional intelligence as the run progressed. When the two vehicles were near perfect alignment in translation and attitude, the center of the despun view appeared as a stable dot. However, translation or rotation misalignment between the two vehicles resulted in a nutating motion of the center dot. At any point during the last 1.0 ft. of closure, if the pilot wanted to stop and make

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x

/ / Wedge-shaped reflected light Ipattern

- DTarget vehicle

Figure 15 Vehicle Relative Pitch Error Cue

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a minor correction, it was relatively easy to accomplish. Single translation pulses in + Y or Z were usually all that was required to keep the vehicles-aligned during the final few feet of closure before docking occurred and the latch lamp lighted.

4.8 The advantage of having a two-man crew to effectively accom­ plish the docking task cannot be over-emphasized. In addition to the co-pilot functions described above, much assistance in scope interpretation and target dynamics was provided by the highly competent personnel who served as co-pilots. This study included an evaluation of the final phase of the approach and the docking task. Follow-on studies during the design definition phase will include acquisition, approach, docking, despin, orientation, separation, and fly-around and will probably require a systems status engineer in addition to the pilot and co-pilot.

4.9 Although this feasibility study did not include the degrading orbital dynamic effects, the crews were able to repeatedly demon­ strate total misdistances (summation of Y and Z errors) of approxi­ mately one-half inch.

4.10 All console displays used during the study were mounted so asto present measured parameter information to the pilot on a vertical scale. The sign convention on the meters afforded the pilot the highly desirable "fly-to" control convention. However, presenting real-world horizontal information, such as Y translation on a ver­ tical scale was not considered to be optimum. For follow-on studies a console layout such as is shown in Figures 16 and 17 would probably enhance the effectiveness of the data being presented and increase the efficiency of crew performance.

4-11 Docking with a 220 sun angle was more difficult as sighting along the 9 and 12-O'clock edges of the ATS-V resulted in a significantly degraded cue as to yaw and pitch attitude error. The reflected light wedge reference was also degraded, but was usable to a lesser degree at a distance of about 10 to 12 feet. When the vehicles were closer together or farther apart, the light wedge reference varied in intensity and definition to the extent that its usefulness was marginal. With these degraded visual cues, when exceptional care was exercised, aligning the vehicles was still possible by utilizing the pitch and yaw attitude error meter and cross-checking the co-pilot's despun view. Under this very limited evaluation, docking was satisfactorily accomplished.

* C-54 SD 71-286 1. Pilot's Monitor 2. Pilot Displays and Controls (See detail on Figure 17) 3. Co-Pilot's Controls 4. Rotation Hand Control 5. Translation Hand Control 6. Co-Pilot's Monitor 7. Indexing Pulse Display 8. Engineer's Monitor 9. Engineer's Control Panel

-0

*00

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0

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Figure 16 Reomended hee-Ma Console Space Division North American Rockwell

-Target radius at 10 ft distance

Target radius at 50 ft distance

HORIZONTAL RATE INERTIAL ROLL ATT

__1 y -. 11[1100 -10

VERTICAL RATE RANGE RATE PIT( H ERR RATE -. 1 DOCKED 1- 18.0- THC MODE DISPLAY 8. M.IN IMP ON

5 ACCEL CMD OFF

+.1 -

Figure 17. Pilot's Monitor and Controls C-56 SD 71-286 Space Division Q11NorthAmerican Rockwell

4.12 Since attitude control in the RMU was normally accomplished by CMG's, the importance of RCS quad locations might not be thoroughly evaluated. However, the RCS will be activated during all translation maneuvers, if a CMG gimbal angle excursion causes gyro saturation, or when the ATS-V despin process commences. The simulated RMU RCS quads were not located on the true Y and Z axes. Previous experience with Apollo vehicle simulations has shown that service module RCS quads located seven degrees off the Y and Z axes have'resulted in undesirable coupling moments being generated during Failures and/or when operating in a degraded mode.

5.0 CONCLUSIONS AND RECOMMENDATIONS

5.1 Under the conditions of the study, docking with the ATS-V was feasible. (PR Paragraph 4.7, 4.9 and 4.11)

5.2 A three-man crew is recommended for accomplishing the docking task with the production vehicle. (PR Paragraph 4.1, 4.3, 4.7 and 4.8)

5.3 Follow-on studies should simulate a baseline RMU vehicle with selectable THC outputs of 30 and 10 MS for acceleration command capability; RHC output discrete rate of I .20/sec with one pound RCS thrusters with the quads located on the Y and Z axes and should include orbital dynamics, off-nominal conditions, and failure effects. (PR Paragraph 4.4, 4.6, 4.9 and 4.12)

5.4 It is recommended that future studies utilize the pilot procedures developed during this simulation. (PR Paragraph 4.7)

5.5 A three-man console with displays as shown in Figures 16 and 17 is recommended. (PR Paragraph 4.3, 4.7, 4.8 and 4.10)

5.6 Docking under limited off-nominal conditions was more difficult but feasible. (PR Paragraph 4.11)

C-57 SD 71-286 A91h Space Division W@E North Amencan Rockwell

9.0 CONCLUSIONS AND RECOMMENDATIONS

9.1 CONCLUSIONS

1. RMU docking to the ATS-V, in general, is feasible from a pilot task and accuracy standpoint. There are enough piloting aids, both ground based and peculiar to the ATS, such that a reasonable accuracy in miss distance and attitude alignment is possible. The resulting docking contact conditions are within ATS geometrical constraints. The limit numbers shown on Table 13 include effects of an off axis spin motion of 1.60 of the ATS-V geometrical center line.

2. The baseline visual system should consist of the lead oxide, non­ smear, TV camera tube providing a minimum of 400 lines horizontal resolution and 300 lines vertical resolution at the pilot's console. The RMU ring tube lights should be bore sighted along with the RMU camera. The light intensity required in 1/4 scale for the simulator was 200 ft-candles. This requirement should be established in a full-scale lighting test.

The 700 field of view, 10 mm lens should be focused on the ATS docking ring at a distance of 5 feet between docking planes. The field of view would be improve format which provides 700 field on each side. A 700 field of view is required to encompass the ATS cannon plugs when the RMU is in the docked position to aid in visual or backup mode latch indexing. A fixed stop of 1.8 is adequate if an adjustable tube trigger voltage can be provided along with an adjustable white clipper circuit to prevent tube burn.

A sun shroud is required around the docking system structure to prevent sun glare from overpowering the illumination of the ATS being provided by the RMU lights. Without the shroud, the RMU light intensity must be increased and the white clipper circuit strength required to protect the camera tends to obliterate image detail.

The derotation optics used in the simulation was adequate for synchronization of the spinning docking system but not for indexing of the latches. Indexing is best accomplished with processed discretes representing a given roll attitude of the ATS and of the PRECEDING PAGE BLNK NOT FILMI C-59 SD 71-286 TABLE 13. RMU DOCKING SYSTEM DESIGN CRITERIA

MIS,DISTANCE LATERAL VEL. CLOSING VEL. ANGULAR MISS ANGULAR RATE PROPELLANT MANEUVER TIME NUMBER (I,4CHES) (INCHES/SEC) (INCHES/SEC) (DEGREES) (DEGREES/SEC) (LBS) (SEC) SAMPLES MEAN LIMIT MEAN LIMIT MEAN Mx MEAN LIMIT MEAN LIMIT MEAN MAX MEAN MAX

.344 1.94 .024 .236 1.35 2.64 .277 2.997 .00094 .0027 .5187 1.45 549.7 838.5 30

oMean = Absolute Value Average of Worst of Two Axis. al0 Limit = Envelope of pilot assuracy encompassing 95.45% of all docking attempts. Miss distance limit includes an addi­ tional 0.46 inches to provide for ATS-V wobble induced offset; angular miss limit includes an additional 1.60 to provide for ATS-V wobble induced misalignment.

Max = Highest value seen of a statistically skewed parameter.

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D0 SID 71-28 Q Space Division NorthAmerican Rockwell

docking system which, when superpositioned, properly orient the latches with respect to the ATS. Latch indexing using the despun image is difficult since the ATS Aft Bulkhead equipment cannot be seen. The co-pilot must rely on memory of external ATS features for proper iridexing. The external features are out of the field of view during translation to dock.

3. The dual purpose gyro control system provided adequate attitude control for RMU docking provided the gyro gimbals did not saturate after attitude alignment was adhieved. The requirement for gimbal status indicators was obvious but the procedure to insure against saturation was not developed. The rotational rate command preferred by the pilots and verified statistically is 1.00/sec in all axes using the Apollo block I hand controller.

The translational control system required to accurately dock the RMU was a .155 ft/sec 2 acceleration command through an Apollo block II hand controller. A switch, to provide a small duration minimum impulse, was required to fine vernier translational rates during final closure. Two pilots preferred .03 seconds minimum impulse, and one preferred .06 seconds. Statistical data slightly favors .06 seconds. A minimum impulse between .04 and .06 seconds is recommended.

A vernier control, for accurate spin synchronization, of + 3 RPM and + 1.5 RPM is required in parallel with and independent of the 2ocking system spinup discrete and derotation optics discrete, respectively. The vernier was supplied by two 2700 control knobs in the simulation. The ability to rotate the docking latches out of the way of the ATS image, prior to docking system spin-up, is required during the angular alignment procedure.

4. The primary display used is the TV image which included range stadia at 50, 10 and 5 feet located on the pilot's TV monitor crosshair which is bore sighted with the RMU camera. The ATS vehicle and bore sighted RMU lights provided the rest of the visual cues required for docking. Attitude alignment cues were derived from two image sources (1) sighting along the sides of the ATS and (2) flying to center a wedge shaped reflection of the RMU light ring on the ATS forward skirt flange and aft bulkhead.

Translational alignment cues were supplied by the center hub of the ATS thermal control louvres and by latch concentricity with the ATS ring.

C-61, SD 71-286 @ Space Division North American Rockwell

Other cues used primarily as backup indicators were the pitch/yaw error meter and the pseudo translational rate meters. Runs without meters show little loss of docking accuracy after considerable pilot training in the use of the TV image. A roll indicator is required to quickly re-establish TV communications in the event that the RMU antenna lobe is inadvertantly rotated off alignment with the ground station.

5. Suitable piloting procedures were developed during the study. Of basic importance is the requirement of a co-pilot to provide the spin-up, synchronization and latch index functions. This permits the pilot to concentrate on maintaining translational alignment accuracy. The sequence of piloting procedures is also of basic importance since at no time is the pilot required to command both attitude and translation simultaneously. This is additionally for­ tunate since the jet logic used prohibits simultaneous translational and rotational commands by giving priority to rotational commands. Since the procedures permit alternate concentration between rotation and translation control, further splitting of piloting tasks is unnecessary though use of a "flight engineer" to monitor space­ craft status might prove advantageous.

6. The effects of a 220 sunlight angle illuminating the docking end of the ATS caused loss of the wedge shaped RMU light reflection cue for attitude alignment. Adjustment of the TV camera target voltage and white clipper circuit was required to reduce glare off of the face of the ATS skirt flange. The glare obliterated the image of the edge of the forward skirt and degraded alignment with the sides of the ATS. No noticeable improvement in the ability to see the ATS aft bulkhead equipment was realized from a 220 sun angle. Docking at a time when the sun angle is normal to the ATS spin axis appears more favorable. The effects of open and closed ATS thermal control louvres were negligible.

9.2 RECOMMENDATIONS

I. The simulation concludes that, with all systems operating, the RMU can be flown with the required precision to dock with the ATS. It is recommended that certain of the more probable system failures be incorporated into the simulation such as single jet failed off or thrust degradation, DPG wheel failure, and high DPG drift rates. Contingency procedures should be developed to handle these failure 'modes.

C-62 SD 71-286 Ad Space Division 'VT North American Rockwell

2. Pilot training and the development of flying procedures from visual acquisition at terminal rendezvous through docking, despin, separation and station keeping on a training simulator­ is recommended to aid in both development of mission success, and in the design of the pilot console and ancillary ground control requirements.

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APPENDIX D

POST CONTACT DOCKING ANALYSIS

SD 71-286 Adlk Space Division IQ NorthAmerican Rockwell

CONTENTS

SECTION PAGE

1.0 RMU DOCKING LOADS AND DYNAMICS...... 1. 1 Maximum Docking Loads...... 1.2 Capture, Stability and Effects of Heat Pipes ...... 1.3 Math Model Description ...... 3 1.4 Docking Conditions ...... 4 1.5 Docking Condition Results...... 7

2.0 DOCKED AND DESPIN SIMULATION ...... 31 2.1 Heat Pipe Model ...... 31 2.2 Checkout of Heat Pipe Model ...... 31 2.3 ATS V/RMU Control System On No Heat Pipes ...... 33 2.4 ATS V/RMU Control System On and Heat Pipes Active . . 34 2.5 Conclusions ...... 34

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D-iii

SD71-286 Ab Space Division NFh American Rockwell

1.0 RMU DOCKING LOADS & DYNAMICS

RMU docking system latching loads and capture stability, both including the effects of RMU control systems, ATS off-axis spin, and ATS heat pipes, have been analyzed in a digital computer study and found acceptable. The digital program is a modification of an Apollo docking program which permits 6 DOF analysis of docking of two spinning or non­ spinning bodies, a spinning docking interface, and docking latches.

The worst combinations of pilot generated docking contact conditions (see Appendix C) were used as initial conditions in the study of docking loads and dynamics.

The following section describes the maximum docking loads to be used as preliminary design limit loads. Capture stability and effects of heat pipes are discussed. The math model and docking contact conditions used in the study are described. The remaining paragraphs present the numerical data describing acceptable RMU docking loads and dynamics.

1.1 MAXIMUM DOCKING LOADS

Table 1-1 tabulates the maximum loads generated during the study. They are of the order considered in the design of the RMU and therefore, acceptable. Additionally, parametric analysis in the next phase should optimize damping and spring constants and reduce the maximum values shown. Further, the maximum values are transients as indi­ cated in the CRT printout in Section 1.5.

1.2 CAPTURE STABILITY AND EFFECTS OF HEAT PIPES

When an RMU latch trigger contacts the ATS apogee motor thrust ring all three latches are released to grip the ring.

The latches described in the math model have a 4 inch distance between grips and a depth to the latch trigger of 2.01 inches. The 4 inch distance between grips permits a 2 inch miss distance, and the 2.01 inch depth to the trigger permits 30 misalignment. Therefore, if contact is within 2 inches and 3 degrees of alignment, capture is always successful. Maintaining capture could be unsuccessful if the loads generated during alignment, and prior to rigidlizing, become too high. -The math model latch is configured such that a radial load'a excess of 1500 lbs. will cause the latch to release the ATS ring. To date for the docking conditions examined in the program the latches have not failed to capture or maintain capture.

D-1 SD 71-286 TABLE 1-1 MAXIMUM LOADS

FX FY FZ MX MY MZ LBS LBS LBS FT-LBS FT-LBS FT-LBS

Case A -1200,+600 0 0 0 0 0

CsaeA -980,+600 -83,+89 -60,69 +2 F26o,- oo +iooo,-1200

Case 3A -1000,+600 -83,+120 -105F+130 +2 -700,+625 -1300,+700

Case 1B -1150,+600 -85,72 -82,+82 +2 -950,+750 -700,+750.

Case 2B -1000,+600 -90,+60 -90,+75 +2 -800,-880 -1000,+1400

N) Case 3B -1150,600 -120,+120 -115,7140- +2 -800,+600 -1300,12001

Case 4B 1050,+650 +120,-170 -100,+125 +2 -600,+600 -925,-700

Case 5B -7000,+600 -731+74 -93,+94 +2 -900,+700 -900,+90Q 1 Case 6B -1030,+600 -85,+60 -75,+83 +2 +850,-800 +1400,-1000

00Case7B -60,+650j +120,-O5 +95,-120 +2 +700,-550 +700,-925 00'

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The math model of the ATS incorporates a description of the ATS heat pipes, which has been verified as discussed in Section 2.0. The fluid in the heat pipes is capable of generating radial loads due to centrifigal force, viscous forces as a function of radial force and an impulse if the fluid reaches the ends of the pipes. In general these forces are less than 9 lbs radially with moments less than 10 ft-lbs.

Since docking loads are on the order of 60 to 120 lbs the affects of heat pipes on capture and rigidizing are negligible. Further, since rigidizing was accomplished and despin started in a fraction of a second after initial impact and each run cdntinued only for I second after initial impact. The destabilizing effects of the heat pipes were not significant during the docking dynamics study. During the 400 seconds of despin the effects of heat pipes were analyzed on a control systems, rigid-body, dual spin dynamics program and the results are presented in Section 2.0.

In the docking dynamics runs, a worst case docking loads condition was run with and without heat pipes. The effect is negligible as shown in Section 1.5.

1.3 MATH MODEL DESCRIPTION

A digital math model of the RMU vehicle, a spinning docking structure, three spring damper latches, and the spinning ATS vehicle was designed and programmed using a large portion of an existing Apollo docking program.

The RMU vehicle was treated as a non-spinning vehicle with an attitude hold control system (dual purpose gyros). A mass-less spinning docking mechanism, capable of applying 2.0 ft-lbs of despin torque .2 seconds after all latches complete drawdown, was attached by a force and moment matrix to the front of the RMU. The spinning dock­ ing system was assumed to be synchronized with the ATS roll rate.

The ATS-V was treated as a vehicle without a control and with cross products of inertia sized such that a stable spin rate was maintained about a spin axis that is 1.60 off of the geometrical center line of its longitudinal axis. A ring was mathematically described to represent the ATS apogee motor thrust ring used as the target docking structure. The digital program assumes this ring to be rigid such that all spring compli­ ance is placed on the RMU side of the docking interface. A simplified model of the ATS heat pipe fluid motion was incorporated to feed viscous damping and end of pipe impulses to the ATS.

Three spring-damped latches are described at three points 1200 apart on the face of the spinning docking system. Each latch is made up of (1) an axial hard point spring representing the spinning docking structure axial compliance, (2) a trigger plane which causes attachment of the latches to the ATS when the ATS ring touches (mathematically) any one of the triggers, (3) a lateral spring that can operate some delta time after the

D-3 SD 71-286 @d&Space Division North American Rockwell

trigger is touched, (4) a (velocity)2 damper attached in parallel with the lateral spring, (5) a second lateral spring that describes load when the first lateral spring bottoms, (6) a preloaded draw-down spring that applies load only when the latch hooks come in contact with the ATS ring flange, and, finally, (7) a (velocity)2 damper that attenuates the draw-down spring and slows drawdown.

A fixed time step size integration package integrates both translational and rotational accelerations and velocities to derive relative positions of the two vehicles and their docking interfaces. As the docking structures impact spring deflections generate loads proportional to the deflection magnitudes for as long as the structures remain in contact. These loads and their moments are applied equal and opposite to each vehicle in its equations of motion until spring forces across the docking interfaces are in static equilibrium. The program terminates on a number of conditions such as time limits, latch miss, unlatch, or zero ATS spin rate.

Basic input data describing the RMU and ATS is shown on Figure 1-1.

1.4 DOCKING CONDITIONS

The docking contact conditions used in the RMU docking loads and dynamics math model are the limit values presented as docking contact criteria in Ref. 1. These conditions are assumed to exist at initial contact:

I Closing Velocity: 2.6 inches/sec, or .217 fps BTW C.G.S.

Radial (Lateral Velocity): 2.4 inches/sec, or .02 fps BTW C.G.S.

Radial Miss Distance* 2.0 inches or .166 ft. BTW Docking. Plane Centers

Angular Velocity: .003 0/s RMU axes relative to ATS Axes

Angular Misalignment: * 3.0 deg. betw. vehicle

* lncludeseffects of 1.60 ATS off-axis spin; & .46 inch miss

The contact conditions shown on Table 1-2 have been analyzed for loads and dynamics during RMU docking.

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5.8'M = 15.5 slugs

= 0.0 1 25' Ixx = 49.0 slug ft Ixy .. ... X Iyy = 44.5 slug ft Iyz = 0.0 I = 43.8 slug ft I = 0.0

SK =500 lbs/in 15" jtoc

-- ] .04' .334' -fjj.04'

K 50,000 lbs/ft

--.168'-

RMU Geometry

I M =31.0 slugs 76.5 RPM I = 98.6 slug ft2 1.25' Iyy = 85.0 slug ft2 2 -- - I-z = 82.6 slug ft IZ

QATS-V Geometry

Figure 1-1 Vehicle and Docking Interface Description D-5 SD 71-286 TABLE 1-2 DOCKING CONDITIONS FOR RMU LOADS

A. ATS Spin @73 rpm Without Wobble

Closing Lateral RMU ATS Miss Distance (Ft) Velocity Velocity Attitude Attitude wrt (fps) (fps) (deg) - (deg) wrt ATS Spin Axis Case ]IA Straight-in .217 0 0 0 0 0 Case 2A 3' Misalignment .217 0 -30 0 0 0 Case 3A 2" Miss Distance .217 0 0 0 -. 166 -. 166

i B. ATS Spin @73 rpm With Wobble Case 1B Straight-in .217 0 0 -1.6 -. 0384 0 Case 2B Max Misqlignment .217 0 1.4 -1.6 -. 0384 0 Case 3B 2" Miss Distance .217 0 P -1.6 -. 166 -. 128 N Case 4B

IO£ Misalign.+Miss Dis. .217 0 1.4 -1.6 -. 166 -. 128 ao Case 5B Lateral Velocity .217 .02 0 -1.6 -. 0384 0 z U0 Case 6B D Lat.Vel.+Misalign .217 .02 1.4 -1.6 -. 0384 0 CD Case 7B

Dist.+Misalign .217 .02 1.4 -1.6 - -. 166 -. 128 o0

SD 71-286 @ Space Division North Amencan Rockwell

1.5 DOCKING CONDITION RESULTS

The following data are the results of the RMU docking dynamic digital math model.

Most of the conditions run were used primarily for checkout of the program. All cases were examined for maximum loads. Table 1-1 presents a summary of the peak loads, and moments occurring in each case studied.

In All cases, capture was successfully maintained. Only case 7B was continued for a sufficient length of time, with and without heat pipes, to assure stability after docking.

The data is presented only for case 7B as being typical of other runs. The nomen­ clature used on the graphs is defined as follows:

NOMENCLATURE DEFINITION UNITS

FSUMAZ RMU Z Axis Force Lbs FSUMAY RMU Y Axis Force Lbs FSUMAX RMU X Axis Force Lbs FSUMTZ ATS Z Axis Force Lbs FSUMTY ATS Y Axis Force Lbs FSUMTX ATS X Axis Force Lbs TSUMZA Docking System Moment About RMU Ft-Lbs Z Axis TSUMYA Docking System Moment About RMU Ft-Lbs Y Axis TSUMXA Docking System Moment About RMU Ft-Lbs X Axis TSUMZT ATS Ring Moment About ATS Z Axis Ft-Lbs TSUMYT ATS Ring Moment About ATS Y Axis Ft-Lbs TSUMXT ATS Ring Moment About ATS X Axis Ft-Lbs RIZX Relative X Distance BTW Vehicle C.G.S. Ft R!ZXD Relative X Velocity BTW Vehicle C.G.S. Ft/Sec 2 RIZXDD Relative X Acceleration BTW Vehicle Ft/Sec C.G.S. RIZY Relative Y Distance BTW Vehicle C.G.S. Ft RIZYD Relative Y Velocity BTW Vehicle C.G.S. Ft/Sec 2 RIZYDD Relative Y Acceleration BTW Vehicle Ft/Sec C.G.S. RIZZ Relative Z Distance BTW Vehicle C.G.S. Ft RIZZD Relative Z Velocity BTW Vehicle C.G.S. Ft/Sec

D-7 SD 71-286 @h Space Division North Amencan Rockwell

NOMENCLATURE DEFINITION UNITS

2 RIZZDD Relative Z Acceleration BTW Vehicle Ft/Sec C.G.S. ZT ATS Z Inertial Position Change Ft YT ATS Y Inertial Position Change Ft XT ATS X Inertial Position Change Ft PST ATS Inertial Yaw Attitude Degrees THT ATS Inertial Pitch Attitude Degrees PHT ATS Inertial Roll Attitude Degrees TCMGZ Control Moment & RCS Torque Z Axis Ft-Lbs TCMGY Control Moment & RCS Torque Y Axis Ft-Lbs TCMGX Control Moment & RCS Torque X Axis Ft-Lbs

Case 1A, "Straight-In" Contact Without Wobble

A docking condition in which all three latches simultaneously grab the ATS and draw down provides the maximum relative closing velocity at impact. Any misalignment permits the individual latch draw down forces applied to the ATS to be staggered in application time, and in some cases one latch may not catch up with the ATS ring prior to contact. Thus, the "straight-in" condition without wobble provides the maximum relative closing velocity at impact.

At triggering, the closing velocity of .217 FPS is accelerated by a 600 lb. tension load to a velocity of 1.2 FPS at impact. The rebound load is 1200 lbs in compression. A tension load of at least 600 lbs is required to prevent the RMU from separating when the ATS goes into a flat spin. Reduction of the impact velocity can be attempted by slowing the latch down such that impact occurs at the same time the latches catch up with the ATS ring.

Case 3B, 2" Miss Distance With ATS Wobble

This condition provides maximum lateral loads and moments on the RMU docking cage. Other conditions that may duplicate these loads are slight variations in cases 4B and 7B. Since the two vehicles do not start out in contact the conditions that exist when contact is finally made are slightly different than those shown on Table 1-2.

Case 7B, 2" Miss Distance, 30 Misalignment, and .02 FPS Lateral Velocity

This case should have produced maximum lateral loads. Slight variations in this condition are planned at a later date. A relatively high axial load of 1600 lbs appears to occur in both versions of case 7B at a poinf in the vehicle dynamics where the two vehicles have first been forced into angular alignment. The load is transient and can probably be reduced through optimizing spring and damping constants.

D-8 SD 71-286 Ad& Space-Division North American Rockwell

Figures 1-2 through 1-11 present the loads and vehicle dynamics generated by case 7B for 6 seconds without heat pipes. Figure 1-4 and 1-5 TSUMAX and TSUMTX shows 2 ft-lbs despin torque being applied at .4 seconds to each vehicle. Figure 1-11 shows TCMGX (control moment gyros) trying to counter torque the vehicle. The run was actually 60 seconds long, and shows the RMU remaining stable and in the process of despinning the ATS.

Figures 1-12 through 1-21 present the same run with the effects of the heat pipe model. The primary effect is to cause a higher frequency oscillation to appear in the lateral loads shown on Figures 1-12 and 1-13. Very little difference in ATS attitude history is indicated on Figures 1-20. The run stabilizes, simular to the case 7B without heat pipes, and continues to despin the ATS.

D-9 SD 71-286 0627-01I-SJ 0012 0000

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0-10 SD 71-286 ?7-DO 1-u

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0-13 SD 71-286 0627-01- J

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0-20 SD 71-286 Uc 7-0 ViJ

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D-26 SD: 71-286 627.01-J 0018 0000 7''

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D-27 SD 71-286 00190627-01-Ji 0000

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D-28 SD 71- 286 0627-01-i8 0027 0000 T t .l i

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2.0 DOCKED AND DESPIN SIMULATION

A C.S.M.P. digital simulation study was conducted with the following objectives.

1. Verify a heat pipe math model by comparison with the known ATS-V data.

2. Use this heat pipe model in a dual spinning vehicle model with the RMU control system to verify that despin can be accomplished.

To achieve these ends a four heat pipe model was incorporated into an existing program for a dual spinning statically and dynamically unbalanced vehicle, and the RMU control system was added to the body without heat pipes. The program without the control system was suitable for checking out the known behaviour of the ATS-V. For the coupled RMU/ATS-V, the program has no provision to investigate flexible coupling between the bodies thus, the high frequency structural dynamic response at docking was not simulated nor dynamic loads evaluated. This was accomplished in a separate digital study as discussed in Section 1.0.

2.1 HEAT PIPE MODEL

Heat pipe dynamics have ...... as four circular hollow rings with a blockage, and a particle sliding inside the ring. A viscous force, proportional to the particle angular velocity with respect to the ring and the normal force, acts until impact with the blockage, and the impacts are assumed to be inelastic. The only real control that can be exercised on the model is to vary the damping coefficient, although the model could be changed to allow some elasticity on impact. Fortunately, this was not necessary since excellent agreement with experimental data has been obtained with the basic model.

2.2 CHECKOUT OF HEAT PIPE MODEL

The program outputs have been compared to the known experimental results shown in the table below.

PRECEDING PAGE BLANK NOT FIOW

D-31 SD 71-286 @ Space Division North American Rockwell

Spin Rate Inertia Ratio Time Constant Configuration Rad/Sec Spin/Transverse Seconds

Transfer Orbit 10.48 0.68 120 Divergence

After Motor Boost 10.28 0.92 11 Divergence

After Motor Ejection 6.8 1. 16 19 Convergence

The stable case energy dissipation was found to occur mostly due to the viscous forces and consequently the appropriate time constant could be easily obtained. A damping coefficient of 0.0035 was found to give excellent correlation. That is

Fvi s = -..0035,Fn

where 0 is the angular velocity and Fn is the magnitude of the normal force.

Energy dissipation in the very unstable case occurs primarily due to impact, and in the mild divergence case due to a combination of both viscous and impact. For both cases excellent correlation has beeh obtained.

Some sample results from the simulation are shown in Figure 2-1 thru 2-3. For all cases the damping coefficient was 0.0035.

Fig. 2-1 shows the results from case C which is the stable configuration. The angular velocity components shown in Fig. 2-1 show the vehicle to be converging to steady state about an axis between the X and Y axes. The time constant is approximately 20 seconds and the steady Y axis rate appears to be approximately -0.23 radians/sec.

This is exactly to be expected since the particles in the heat pipe model are converging on the positions to maximize the spin axis moment of inertia, i.e., the upper particles are converging to 3 7/2 radians and the lower to r/2 radians. This can be seen by examination of Fig.2-1 (c) where LI and L3 are the angular position in .radians of the particles above the c.g. and L2 and L4 the angular positions of the particles below the c.g. This position of the particles gives a major axis tilt 1.9 degrees away from the vehicle X axis. The limits on the particle travel are 0.67 and 5.61 radians respectively from Fig. 2-1(c) shows that impact occurs only on particles 2 and 4 at approximately 2.4 and 2.1 seconds, respectively.

D-32 SD 71-286 Space Division ONi NorthAmencan Rockwell

Fig.2-1 (b) is included to show that very little motion between the two bodies occurs and consequently the use of a dual spin program for simple spinning body is quite valid.

Fig. 2-2 shows the results for case B, the very unstable case. Examination of Fig.2-2(a) shows that the divergence is more linear than exponential, and time constants as low as 4 seconds could be implied. However, for intermediate values a time constant of 10 seconds looks reasonable. Examination of Fig. 2-2(c) shows substantial impacting all on the lower limits. Examination of Fig. 2-2(b) shows that the relative rate between the bodies does become significant, but by that time the transverse rate is large compared to the spin rate and consequently impacting is to be expected.

Figure 3 shows the results for case A, mild divergence, and from Fig.2-3(a) a time constant of 140 seconds is apparent. Examination of Fig.2-3(c) shows that particle 4 (the furthest forward) particle impacts a great deal on its lower limit but none of the other particles impact at all.

The above results show that the heat pipe model used gives extremely good correlation with the observed data and can be used with confidence in the ATS V Despin maneuver.

2.2.1 ATS V/RMU With Heat Pipes No Control System

Hand calculations have indicated that the time constant for the ATS V/RMU dual spinning vehicle should be several thousand seconds. To check this a simulation run was made and the results are shown in Fig.2-'4. Fig.2-4(a) and (b) show that the divergence in 100 seconds is virtually undetectable, which would be expected with a very large time constant. Two additional cases are included in Fig.2-4 compared to preceding cases they are (b) and (e). The body 2 (ATS V) rates are shown in (b) and the heat pipe particle velocity in (e). Case (e) shows that most of the particle action is a low amplitude, high (ATS V) frequency oscillation. Superimposed upon this is a slow drift as the particles try to move towards a new stable major axis spin orientation. Only particle 3 hits a blockage in this case, as indicated in Figs.2-4(d) and (e).. For all the ATS V/RMU cases the particles are numbered going forward on the vehicle.

2.3 ATS V/RMU CONTROL SYSTEM ON-NO HEAT PIPES

A sample of the control system performance is shown in Fig. 2-5. The case was chosen primarily to evaluate the D.P.G. performance capability when saturation does not occur. Gimbal angles of less than 0.55 radians can be observed in Fig.2-5(b), and no R.C.S. torques can be observed in Fig. 2-5(c). The time constant of 0.26 seconds can be obtained from Fig. 2 -5(a) as expected. This run verifies that the D.P.G. control system is performing as anticipated.

D-33 SD 71-286 9 Space Division I1% North American Rockwell

2.4 ATS V/RMU - CONTROL SYSTEM ON AND HEAT PIPES ACTIVE

Preceding sections indicated that the affect of the heat pipes on the ATS V/RMU should be negligible and that the control system is performing as calculated in the D.P.G. non-saturated mode, In this section the affect of R.C.S. firing on system per­ formance is evaluated, and a simulation of a complete despin is described. Figure 2-6idnd 2-7compare the vehicle and heat pipe response with and without the R.C.S. for a no despin case. The initial transverse rate of -0.2 rads/sec was chosen to ensure gyro saturation. Comparison of Fig.2-6(a) and 7(a) shows that performance is better with the R.C.S. jets off. The angular rate decays to 0.14 rads/sec with the jets on and 0.12 with the jets off.

This is primarily due to the control system lags. The R.C.S. can be made to help the damping by introducing some lead into the system or changing the control system logic from a simple gyro desaturation scheme. However the performance looks acceptable with no changes at all. Comparison of Fig. 2-6(d) and 7(c) show that the heat pipe particles behave essentially the same whether or not the R.C.S. is active.

Section 2.2 stated that the ATS V should be spinning with a steady transverse rate of 0.23 radias/sec. in the reference axes, because in the analysis the principal axis was tilted by unsymmetrical location of heat pipe fluid particles. Immediately after impact if no disturbances were introduced this rate would drop to about 0.04 rads/sec due to the increased transverse moment of inertia. The results from a worst case docking simulation study (Section 1.0) indicate that after the structural oscillation dies down the transverse rate would be approximately 2 degs/sec. as expected. Since this rate is too low to overexercise the control-system, initial conditions for the despin simulation were chosen to be wyo = -0.04 rads/sec and wzo = 0.1 rads/sec. The results of the despin simulation are shown in Fig.2-8. A print interval of 0.5 seconds was used to include the points on one curve. The high frequency (ATS V) oscillation has a period of approximately one second so that the print interval was somewhat large. However, the results show that despin is accomplished very well even with the R.C.S. on. Examination of Fig. 2-8(e) and (f) show that all the particlei impact and stay at their upper limits which is to be expected although a little chattering of the most forward particles is evident.

2.5 CONCLUSIONS

This simulation study has demonstrated the following points.

(a) The heat pipe math model postulated gives a very good simulation of that existing on the ATS V.

(b) The effect of energy dissipation due to heat pipes on the ATS V/ RMU stability will be negligible.

(c) The control system with R.C.S. OFF is better than the R.C.S. ON

D-34 SD 71-286 @ Space Division North American Rockwell

(d) Even with the R.C.S. on the control system performance during despin is excellent.

(e) This performance can be improved by either switching the pitch and yaw R.C.S. control off, or modifying the control system logic to make the R.C.S. reinforce rather than oppose the D.P.G.

* D-35 SD 71-286 REFERENE FRArE ANGULAR VELOCITY CCOPENTS RADIANS PER SECct 0001 0000

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0 2 4 6 8 10 12 14 is is 20 22 Lx 26 -IUR-AS 2-1 - C T- FE MTREETO TIME

1t0 12 14 16 16 20 22 24$

-SD 71-28 FIUR -1 A E C AT-V AFERMOO EJCTO -. 01------

SD7-8 LD-37 PARTICLE POSITION 0006 0000 - = --: - 'b ; I - -'t .L . -J ,_ -_"---­,_,- ,­ -97J

3 1 II I I:_ rI ,! I ! ' 'Ll l! 0 2 4 6 8 10 12 14 is 18 20 22 24 26 TIME

I S :-V -

o 2 4 6 8 10 12 14 16 18 20 22 24 26 TIME I I ......

4 -1 ---­ ITj6! __ _ T~-

-­ j- r =1=-- -A

o 2 4 6 8 10 12 14 16 18 20 22 04 as TIME

FIGURE 2-1c CASE C ATS-V - AFTER MOTOR EJECTION

D-38 S[5 71-286 REFERECE FRAME ANOAR VELOCITY COTMCtENrS RADIANS PER SECOND 0008 0000 0000 25 ii ___i__,__

f o ______t5 ______LO-I-- *----1--- i- I

- -- - -i ­

2 I I :z -ii - Tj' L2 4 _ ] t 2 141 2 - 2_2

TIME 4 -- - I I 11111111 olIf I I-'---zz:

2 I i I// ' i II A I I [ I ,

0 4 6 1t 12 14 16 IS 20 22 24 26 - i I I TIlE - , , i 4l - -. -

I I ______]______1-

O 0 12 14 16 18 20 22 24 26 'I ME TIME 0 4 6 10 1 2 4 is1s 0 -2-a D-3

SDI71-286 2-~~~~~~S77HI;Ttt8 71___ ZZ6 RELATIVE MOTION OF SECONCARY VEHICLE RADS PER SEC . RADS 0011 0000 "0­ 0 ______-__ "F-______

T-: -

I I ± 1tt II II I-I I I 1 1

-J.I -

0 2 6 10 121 14 1 Is 20 22 24 26 TIME

- -y------iH i--­ z± 4 -,

-1L 11

- + E - ±3-4 J II II I I

TIM

*4...... __ _...... LT..._ - ~'4k

0 2 4 r a 10 12 14 1S is 20 22 24 25

FIGURE 2-2b CASE BATS-V - AFTER MOTOR BOOST D-4 0 SD 71-286 -1-1±±1PARTICLE POSI TINi~ 2i694-9.7Vt:

, , ,II ­ a------Li lb I _-j ;1 ~_ t k2'! i l 4' -!--'TI- 'HiTIM'E ii -i -l zi, .±t.±I I!,

6 tol40 1 14 15 18 20 22 24 2

ElIM

L 4 2 LI #tI_ - I --- I

~*---.-..- ,--'tII I it 0 2 4 6 a to 12 14 16 is 20 22 24 2:5 ------lI I-Li]L. TIME - I------I I :_ 11 o -, I , L 4 0 +r wi

10 12 14 1 18 20 22 24 25 TIME

i \ i i I + I i I .20 10 12 14 is, 22 24 2. 4 1 1 N I 1 I

TIME ;:: _V______i ,---. ______rJ if - _i __II i zI:'t--~ ­ ,2',u "_: ______' 10 Lin... 15 lB ' 11~~S O 12 14 71-21t86 1 8 20 -22 2 25

FIGURE 2-2c CASE BATS-V - AFTER MOTOR BOOST

D-4 1 SD 71-286 REFERENCE FRAE ANGLLAR VELOCITY COWONENTS RADIANS PER SECOND 0001 0000 25

105

oBi 0 to 20 30 40 50 60 70 "80 90 100 TIME .6

.4

-. -: - - - - .~I.B.1 1A III

I a 30 40 50 60 70 80 so 100 TIME

.6

.4 - -C T I I.0 f IIA A :1 i a II t I L J- .L

1 +F I I I~I1 i 1 1 1 111 1 1 +4 : I I Ii -LL __ ,___i-L-11 v

0 10 20 30 40 50 60 70 so-- 0D9 TIME

FIGURE 2-3a CASE A ATS-V - TRANSFER ORBIT

D-42 SD 71-286 RELATIVE MTIONCF'SECMOARY VEHICLE RADS PER SEC . RADS 000'4 0000 .0

-IO-., F 1,

-.004 tt

-. 00

S .00 I.OL

-.016NI

-. 018

0 10 20 30 40 50r 60 70 80 90 100 TIME

.2

-- p - H

I

-.3 If

0 10 20 30 40 50 60 70 80 90 100 TIMEt

FIGURE 2-3b CASE A ATS-V - TRANSFER ORBIT

D-43 SD 71-286 PARTICLE POSITION 0006 0000 3.0

2.0

10 20. 30 40 50 60 70 80 90 too TIME 2.4

2.0

1.6

1.4 0 10 20 30 40 50 so "70 80 so too TIME 5.0

4-01

10 20 30 40 50 60 70 0 90 100 R

6 4.L---W - - - -

o. ------W L­

TIlIE \ ~~n yf\ III

0 - -IK -f III I I 0o 10 30 40 50 60 70 so 90 too TIME

FIGURE 2-3c CASE A ATS-V - TRANSFER ORBIT

D-44 SD 71-286 REFERENCE FRA1IE ANCLLtAR VELOCITY C0IPOENTS RADIANS PER SECN 0003 0000 .000015

-f444-. I 1

II W it I ff__ H 1' IJ-1 1111

-. 000005 0 10 20 30 40 50 60 70 80 90 10 TIME I I06I I I I I 'I !

.0:C2 I I I I II I l l I 1 I I III I I I I I I I I I I I I y | I II I I I II t I I I I II I III iI Iiti III

-. 06

I I I I I I I I I I I I I I I ill I I I A] I I -A

, I I I I I ff| I I I

I 1111 tt 1 1 JII I II Iti 1 1 i II I -41 -l I'I II I I z! It II -. o4i 'b I IllI I { I I llJ i I fill I | I l l 1 ~ = I 00S LlILIl ll , ~ l -.IGURElo ~~~ I AT-VRM NO CONTROL SYSTM, I I

0 to 20 30 40 50 60 70 so 90 O00 TIMdE FIUE24 TSVRUN ONRLSSE

I / III I II I . I5I ImID-4 h .OH~I ~ ~ ~ l IIUII~ ~ ~~~~SII \II\ ilIH 71-286I / ' I BODY TflO RATES 2169-97A

7.6500 ___

2

08 o 0 0 3 404 0

0 10 20 30 40 50 60 70 80 90 IO1 TIME

.04

o 10 20 30 '40 50 60 70 80 9 103

TIlE FIGURE 2-4-b ATS-V/RMU NO CON1tROL SYSTEM 0-46 SD 71-286 RELATIVE MOTINCt OF SECONDARY VEHICIE RAtS PER SEC RADS 0004 0000

> - - 7.-490-_-

~ I j J I 1 1-1 LI75+ I I- I - II

7.64e95 - i_ _-

S 0 -. -­

-0 0- -. 0 -1 -___- -20 30 40 50 60 70 so 0 00 TIME 800 H- - - -±

700

7.007 11i 1:

00

0 10 20 30 40 50 60 70 80 90 100 TIME FIGURE 2-4c ATS-V/RMU NO CONTROL SYSTEM

D-47 SD 71-286 PARTIcaE POSITION 0006 0000

L I.5 2.S '.4

a 10 20 30 40 50 60 70 80 90 100 TIME 1.6

L 1.4

1.2 1

2- -­

0 10 a 30 40 50 50 70 so so 100 TIME 6.0

L 3 ------:

10 20 30 40 50 60 70 s0 90 1 00 TIME 5,0

L 4.51

0 20 3m0 40 50 60 70 s0 90 100 TIME FIGURE 2-4d ATS-V/RMU NO CONTROL SYSTEM

D-48 SD 71-286 2694-71O PAFRTICLE VELOCITY

---- t 7- - --vL1~ A - _ _ _ .0211 0 AAA I - VI " ^ = ' L .02 - I '. .. ,i, vvvv ^fvv',I f., ,_ -I

I0 20 30 40 50 60 70 80 90 too TIME

-o L .

-.Oi

O Io 20 30 40 0 60 70 80 so 10 .05 TIME .15

L o - I , 3nr

A0 I. I I I I I I I

I I I t TIT o01 : :0 : A A 111101 ll If l A*

I iV tt t 111u V 1 I, Ii,,, -. 10 20 30 40 EEIEEso 70 so 90 too

.:on~ ~ IGR 2-ll AT-VRM NO, CONRO SYSTEMn~l- JAAl. "70 o 9 0 oo 0 o 2 0 3 0 4 0 ,50 60

FIGURE 2-4e ATS-V/RMU NO CONTROL SYSTEM

D-49

SD 71-286 REFp-ENCE FA ' A3,lAI ELOCITY C cA N- TS R, S F,n SECDOr) 2c3g0o CI .0003

.0002 -­

. -i -

,XI I I-I I I I

0 it v I II , _i_ __ v I vIr v -,--" - I.Ij 1L1IrIIzI I I tI I I1 -N jryz - =~ I1A

I I

a 2 4 -- I 6 a 10 12 14 16 100L16 20 22 : ,24 , . IS es I 3 TIME I t A:h I I 1\ I1 1 101 1413 1, 1 1 1 1 1 1 I I I I

01-. J: _ <.i Tit' FII I_•1 It I Ij III4 a 2 4 6 a 10 12 14 16 Is 20 22 L4 a6 28 30 TUE .0D

SD 7

-ZI U

I ~ - . - . - - I-I

.1l.\I j II , -­ - i-r- ]i

-.02 -l 1 _ i i H AJ A ]M 04 ' 6 8 10 I2 14l 16 18 20 22+ 24 ,30 28 TIlE[ FIGURE 2-5a ATS-V/RMU - CONTROL SYSTEM ON - NO HEAT PIPES 0-50 SD 71-286 S9-0I-j" e 4 AtlCA A/CSL k OCOHI OC00

._ ,f j±..L: : l±-

S -. 010 -- - -i-_--IL k I I t L

. . . . L-I. .. --...... I , I I Ir I1 I IA

I P1 i i i

-H01___, Iiii':::itI _ I •-- ::1

o 2 4 - a 10 12 14 1e 18 20 22 24, 26 28 30 TMlE Ii nI I I I I .5

s -4 1_v _ III I SI 0 2 4 8 10 ~12 1+ 1 18 20 22 4 28 30 4 ~TIlE

. a I

-T- L .4

A - I. - -

S,- 71

A z

-.4- 111--- O 6 a8 1 12 14 16 Is 20 22 as42 28 30 TIME

FIGURE 2-5b ATS-V/RMU - CONTROL SYSTEM ON - NO HEAT PIPES

D-51 SD 71-286 u es2S94-01-01 0005 Moao 1.0

-~~-L-. - 4 R - - - -2­ 0 -. --- -­ LLL 1I1'

o0.517T I xA :, FZ LJ

0 2 4 B8 0 { 12 14 I16 I is 20 22 24 5 28 :30 TIME 0.5

R 1 - -- ::

R a t10- - -12 -, 14 16 18 20 22 2'4 25 28 30 TIME

I I I I - 0.5 -0.5 1 1 1 I

0 2 4 6 8 10 12 14 15 18 20 22 24 2 22 30

1.0 TIE

0.5

-1.0I I I 1 1 1 1 1 1 1 T - R S: 0

-0.5

o a 4 8 8 10 12 14 IS 18 20 2224 26 29q 30 TIME

FIGURE 2-5c ATS-V/RMU - CONTROL SYSTEM ON - NO HEAT PIPES"

D-5 2 SD 71-286 REFERENE FRAE ANGtLLAR VELOCITY COMPONENTS RADIANS PER SECCO 000126914-99-01 0000 .0 1 __ _ "

A P I ]

.0 - ,1 1 IVA

I1 1 1. I'I --

I / -

II -0- I I -I-­

0 5 IO 15 20 25 30 35 4a 45 50

TIME .2 - 1 1 1 1I

a- - _ '7i'.,- -: I : A I I-I :,3§i~iII :

------I-"ili.itIii:iI IIt1 -: lot ' itA I v -s ­

-. 3

TIME

.2 IM FI I I I11/ 11 - I A i 1 t I I I iti i ]VIZ­ t

11 ilII I I it

-i'44 :,PltIt!,44 itt I

.2 __ _ _ _ 5 I 15 20 25 30 35 40 45 50 TIE

FIGURE 2-6a ATS-V/RMU COMPLETE MODEL - NO DESPIN D-53 SD 71-286 GWALAI-aCES 0005 0000

", I ii,, i - ----:-c _1- -I I -. iJ -- - N R I i ik i

TIM

I I I i S1 I

2 , I , I I -t- I\ / I I\ I I I I/, II I .,. it I I I I I I I I t l - ::-:l/: II

1 1--- 1 11 1 1 _ III I I o i~ftl5/ Iii10 I/ 151I t II20 25 I1 130 1 m5 40 45 50

TIME 0! -- : _ L-t I ! I

:0 I I _ 6 IUI M II -- NO I!

D-54 S 71r28-- II ~~I_I

FIGURE 2-66 ATS-V/RMU COMPLETE MODEL -'NO DESPIN

D-54 SD 71-286 2694-99-0Cl R.C.S. TOiQEs 0006 0000

- L

__r_ __-_ - t , 'z !- _ _ __ -e~1i: - I Iit_ I

-- tzi -'-- - -'Tthii t-

5 tO 1.5 20 25 30 35 40 45 50 TIME 6 VI------I ±jE'

S1' I I I I I - - I eRt - I- - L I IME I

-5I 45 50 a 5 t 15 20 5 30 35 40 _I IhtI1 1I --I zszh AI I14 1 11 3:-L­ z -- - -I __J II II f1-

C ;TI -101

FIGURE 2-6c ATS-V/RMU COMPLETE MODEL - NO DESPIN

D-55

SD 71-286 PARTICLE POSITION 0007 0000 2.4" 2.2 ------

T~-TE

0 5 10 15 20 25 30 35 40 45 50 TIE 1.60

1.4

_ . I 4.50 iTI 4------Lt . !!

o 5 1o 15 20 25 30 35 40 45 50 TIME I Al­

5.0

- , I , . *II. I _ _ __- W

I0 30 35 40 45 50 15 20 25 TIMEI 4.-i -: -­ -_ ' -_ _ _ :4

FIGURE 2-6d_ OPEE OE ODS AT-/M

D-56 SD 71-286 2694-99-01 REFERENCE FRAFE ANLLAR VELOCITY CCtPONENTS RAOIANS PER SECOND 0009 0000

.05 .o , , M| 7 H _ _

.0 - --. --

a 5 10 15 20 25 30 35 40 45 50

.3

:: - I r i:: ---+-+--- - ­ -- .4__ L 14 -4

0 5 10 15 20 25 30 35 40 45 50

TIME

.4

.2j __1lIjI I I

.4 I - 'I Sut II .0~~~Tj -*T->_ 2

0 5 15 20 25 30 35 40 45 50 TIE

FIGURE 2-7a ATSV/RMU - COMPLETE MODEL - RCS OFF - NO DESPIN D-5 7 SD 71-286 2694-01

GIMIBAL ANGLES 0013 0000

itil

I ,i Yc I A rI i I i _,. 5 o!0i I 53T iIB 5 4 55 ITIM

0 5 10 15 20 25 30 35 40 45 so TIME 2.0 _ ~i

/li U\ li --- I

-II - IF

1.5--- i _ a I0 15 20 25 30 35 40 45 50 TIME 2.0

_. -. - -J- ,,, 1 It [-L _\ \ -

I II - I

. _ o.___ I__ H _ 1 -, (V I. I -II I I-I FIGRE -7bAT-/ - COPEEMOE CSOF-NODSI

TIrE FIGURE 2-7b ATS-V/RMU - COMPLETE MODEL - RCS OFF - NO DESPIN

D-58 SD 71-286 PARTICLE POSITIONJ 2690015 -.99-0000 0J 2.2

-H 9 KI------­

1.4- t14iL­ C LI -. J.. 10 15. 20 25 30 35 40 45 50 TIME

1.6 -

I- i

1 I '[ TI

a o 1 20 25 30 35 40 45 50 TIME 6.0

1 _'

4. H_ FIGUR 2-7c__ ATS-V/RM MOE F ODS -OPLT -C 0 0 15 20 TIEL5 30 35 40 45 50O o 5 10 15 20 2'53 5 40 45 5 TIlE

D-59 SD 71-286 0001 000 RFERENCE FRAME ANGAn VELOCITY Ca9Gn4TS RADIANS PER SECOND .06

I.04 I It ' II Ill I I .. I I... I." I' I I I I ,

0 204 60 8 100 0140111 1801 fl 180 2fi4l20 20 30l2 10 40 30 3 0 2

.1M' .

iii

.0 I: I

It rI- I21

0 20 40 60 80 100 120 140 10 160 200 220 240 260 2-80 300 320 340 360 400 '420 TIE TIE .0

FIUR 2-8 AT-/M COPLT MODEL - DEP WIT RC ON­

S.02 -28 441D-110 0~ ~~ ~ 10101016~ ~~~~~~~S 0 2020202807041-8 0 2 4 303 0 2 RELATIVE MOTION OF SECONARY VENIC-E RADS PER SEC * RADS 0004 0000

Ii l I lPI~ l l l l l l l l l i I I i } l I~iI l l I I I S

sI II I I I l IIl l t ~ If I I II I I I I lII lI I I

4 -

I I II II II II I I I I I I~:J T I I I fI I I I I I I I I II

" ' "

Il I I I I I 120 ------I I I-itIII Ii% I1 I IIII I I I 14 I I

TIME

locl:= " ll " "I" " " ;.1" " " " " " " tIl - "I II' .

H I I I I I I I

f il lI oi

0 20 40 60 80 100 120 140 160 180 200 220 240 260 280 300 320 340- 380 380 400 420 TIME

FIGURE 2-86 ATS-V/RMU COMPLETE MODEL - DESPIN WITH RCS ON

D-612 SD 71-286 --j GIMBAI-ANGLESGUSAL.A~aES0005 269400004

I+ - 0.4 ----- I I I I IA ll

S :-0.6

:" -q~~~~~31 "lT11- "l#7-[ 1. JT l

-1.FI 0 20 40 60 80 100 120 140 160 180 200 220 240 260 280 300 320 340 360 380 400 420 TIME A 2

S I I I I 1 11ill 1 0 20 40 60 80 100 120 11180 200 20 240 20 280 300 320 340 30 30 400 420 H A z

m . . I40 101OO10 0 14 6] 10 20 ;0 20 20 20 30 30 0

Til

SD 71-286 R.C.S. TORQUES oos 0000 0

...... '...... * 4hLL7 .----p -- --±Lkv !411BI lL .f;+ ..

st A tit+ ,1"M' . i-t. +

-4 .T

0 20 40 60 So 100 120 140 160 180 200 220 240 260 280 300 320 3-40 360 380 400 420 TIME

." [1 I I I

04 0 l01 120 140 160 180 200 220 240 260 280 300 320 340 36 230 40040 TIME -61"'" I T 1 1, I- I I II" " " " " " l l ' " eo 4 o s .160 P,18,, 4,, ,, ,, ,,200 240.,,,, , ,280. , 30,,,-630,,,, --- 0 4 TIMEI

.+SLDJ 71286

5 I t I I I t ; 1II, I I I I1 Io I I I I I II~il I I ------I -- R I !l I ' " ' ' R 4 1 1 1 11+,+ ' '~ I I II---I-- I I I--- I '-!I---- ­ -I- l

0 a 40 60 80 100 120 140 160 180 200 220 240 260 280 3013 32?0 3'40 36 3W 4W0 .0 -- 4 TIME FIG"""U'REil2-d T""/""CMPETIODL S'NWIH'CSO I ~I "'""'"I I +""+"""'"""'"""""'"''""SD! I +I1 1 : I I I I I I I I I I I I I I I II II I I I I I "7"1-286"iiIIDI-II6I I3 PARTICLE POSITION 000 0000 6

00

O 20 40 60 So 1oo 120 140 160 isa 200 220 240 260 280 300 320 340 360 380 400 420 TIME 6

2 7 l! l li l ~I~ t l' l ! t!! ! !l lf ! l~ fi ltl tItII ! l ! l i l i II ll l l l l I I I I ~ lI l I iI I I l l i l ll i l lIII I I I Il IIII l i l l i Il l i lt l l l l l Iti l l V I ~ il II I II I I I I l l 1 f il I IHll l l ' ll I ll l --­ l l l l l o 20 40 60 80 100 120 140 160 180 200 220 2X40 260 280 300 320 340 360 380 400 420 TIME

4.0

4.5­

o; a0 40 60 80- 100 120 140 160 180 200 220 240 260 280 300 320 340 360 380 400 420 TIME

FIGURE 2-8e ATS-V/RMU COMPLETE MODEL - DESPIN WITH RCS ON

D-64 SID 71-286 PARTICLE VLOCITY000 .2

I II I 0i I I i Ili I l t IlIt IIIf i l i

L .1 " " "I I I I I !

Dl l l l lii i it I -iI 0 20 40 60 80 100 120 140 ISO 180 200 220 240 260 280 300 320 340 360 380 400 420

TIME

0 0 40 60 80 100 120 140 160 180 200 220 40 20 20 00 320 0 360 0 400

.15

0 20 40 60 0 100 1;?0 140 160 180 200 220 240 260 280 300 320 340 360 380 400 420 - iI i l lI I I I I I I I 1 1 II I I I I 1it 1

TIME

.101I SD1- 1

0 20 40 60 1O00 120 140 160 180 200 220 E40 260 280 300 320 340 0 80 400 2 TIME

lI I l I1 it itl l D 11,I1 IIIII

ii I I I I I

0 m, 40 60 8 0 10 10lo S 0 aa Lql6l20 30 30l 4 6 0 40 92

TIMIEI

FIUE28 T-/MUCMLT OE SIN IT RCS ON

SD71-286 .b Space Division North Amencan Rockwell

APPENDIX E

TYPICAL PROCUREMENT SPECIFICATION

SD 71-286 $it Space Division 4 W North American Rockwell

6.00 3.50

4.65

.5007Y? b~

TABLE I CONFIGURATIONS

01 I & T CONTROL NUMBER

E ME495-0011-001

O I & T = IDENTIFICATION AND TRACEABILITY E = EXEMPT FROM TRACEABILITY

NOTE: ELECTRICAL TIRMINATIONS AND MOUNTING PROVISIONS (TBD)

0 NR/SD PROCUREMET SPEC: NONE

TOLERANCES, EXCEPTAS NOTED TOLERANCES ON .013 ThFU .o0 : +.00I -. 001 .301 TIeU .7350+.005 -. 001I IG.S.XX- 0103 HOLES NOTED .041 THRU. 310: +.00 -. 00 .7351THUI i.0:+.007 -. OI CODE IDENT NO. ANGLESD NClAALtS .OE'DRLL' .131 TlTHU 2? :.-.001+. I.001 TH U 2.000.+.010 -. 00 03953 .XX-*.0I0 DIL .230 THeU .30 .4D .00

J. WOLTZ SeSPACE DIVISION CONTROL PART NORTH AMMICAN ROCKWELL CmPORATION CHK Y M1f14L SOAXVAAO*W . V i M

a 'INVERTER, THREE-PHASE

SPEC CONTROL DRAWING SET 1 of 6 ­

E-1 Space Division North Amencan Rockwell

1. DESIGN

1.1 GENERAL REQUIREMENTS. THE DESIGN AND CONSTRUCTION OF THE INVERTER SHALL CX'PLY WITH THE APPLICABLE REQUIREMENTS OF SPECIFICATION MC823-0007, DATED 5 NOVEMBER 1970. THE INVERTER SHALL PROVIDE AN OUTPUT-CURRENT/ OUTPUT-VOLTAGE hEULATED, 3-PHASE, 3-WIRE QUASI-SQUARE WAVE OUTPUT USED FOR DRIVING SIX HYSTERESIS SYNCHRONOUS GYRO MOTORS IN PARALEL. IN ADDITION, A 3-WIRE FILTERED VERSION OF THE 3-PHASE QUASI-SQUARE WAVE OUTPUT SHALL ALSO HE PROVIDED FOR LINE-TO-LINE SINE WAVE EXCITATION OF THREE RESOLVERS (ONF PER PHASE). THE INVERTER SHALL 'BE OF STATIC, SOLID STATE DESIGN, USING SILICON SEMICONDUCTORS.

1.2 CONFIGURATION. THE SIZE, ELECTRICAL TERMINATIONS, AND DIMENSIONS OF THE INVERTER SHALL BE AS SHOWN ON SHEET 1.

1.3 WEIGHT. THE WEIGHT OF THE INVERTER SHALL NOT LXCEFJ: 4 POUNDS.

1.h INPUT POWER.

1.4.1 INPUT VOLTAGE RANGE. THE INVERTER SHALL OPERATE WITHIN THE SPECIFIED LIMI WHEN SUPPLIED WITH 22 TO 34 VDC AT THE INVERTER'S INPUT CONNECTOR. THE INVERTER SHALL ALSO BE CAPABLE OF OPERATION AT UNDER/OVER VOLTAGE CONDI- TIONS OF 18 TO 36 VDC PESPECTIVELY AND SHALL NOT INCUR ANY PERMANENT DAMAGE DUE TO SUCH EXPOSURE.

1.4.2 SOURCE IMPEDANCE. THE INPUT POWER SHALL BE SUPPLED FROM A SOURCE HAVING A MAXIMUM IMPEDANCE OF 0.5 OHMS.

1.4.3 INPUT POWER INTERRUPTION. INVERTER TO RECOVER WITHIN 50 MILLISECONDS AFTER CESSATION OF INTERRUPTION.

1.5 ELECTROMAGNETIC INTERFERENCE. THE REQUIREMENTS OF MIL-STD-461, DATED 31 JULY 1967, CLASS IIA EQUIPMENT SHALL APPLY EXCEPT FOR THE CONDUCTED NOBE IN THE OUTPUT LEADS (TO BE SUPPRESSED WITH THE SPECIFIED REFERENCED GYRO LOADS AND PROPERLY SELECTED INTERCONNECTING CONDUCTORS).

1.6 INSULATION RESISTANCE. INSULATION RESISTANCE SHALL EXCEED 100 MEGOHMS AT 100 VDC. 1.7 DIELECTRIC STRENGTH. THE LEAKAGE CURRENT SHALL NOT EXCEED 3 MILLIAMPS AT 70 VAC AT 60 CPS. 2. PERFORMANCE

2.1 VOLTAGE PHASE RELATIONSHIP. PHASE RELATIONSHIP SHALL BE 120 DEGREES PLUS OR MINUS 10 DEGREES WITH A-B-C ROTATION. a

TOLEAANCES, EXCEPT AS NOTED TOt EUACES ON .013 THRU .040: +.001 -. 001 .501 TIRU .750: +.Z -. 001 , OtES NOT .041 THRU. 130: +.00D -. 00! .7SI THU l.0D0-+.*7-.01 COE IIENT NO. ANGLE.,l0.XXANGES3O . '-*.01ECIMLSUt "DRLL"'DRILL 230131 THI(UTHRU .5I0.22?: : +.004+.00 -.001001 1.001 TiC.) 2.DCOW:0l0 -. 6Dl 03953

DRby J. WOLTZ -17- 14 SPACE DIVISION CONTROL PART j NORTH AMERICAN ROCKWELL COPORTION CH(SNY =AXVM * l.CkM APflOVtO BY INVERTER, THREE-PHASE ME495 'O01

SPEC CONTROL DRAWING SET 2 of 6

E-2 Adi l Space Division 9W NorthAmerican Rockwell

2.2 FUNDAMENTAL FREQUENCY. THE FUNDAMENTAL OUTPUT FREQUENCY SHALL BE 400 CYCI.S PER SECOND PLUS OR MINUS I PERCENT. CONTROL OF THIS FREQUENCY SHAll BE SUITABLY ISOLATED FROM THE LOAD SUCH THAT OSCILLATORY VARIATIONS OF LOAD CHARACTERISTICS (I.E., -HUNTING" OF GYRO WHEELS AT SYNCHRONOUS SPEED) SHALL NOT MODULATE THE INVERTER OUTPUT FREQUENCY.

2.3 TYPE OF LOAD. ON THE QUASI-SQUARE WAVE OUTPUTS, THE INVERTER SHALL DRIVE, IN PARALLEL, SIX BALANCED, 3-PHASE, Y-CONNECTED, HYSTERESIS SYNCHRONOUS GYRO MOTORS PER ME467-0022, DATED 5 NOVEMBER 1970. ON EACH PHASE OF THE FILTERED SINE-WAVE OUTPUTS THE INVERTER SHALL DRIVE ONE LINE-TO-LINE CONNFCTED RESOLVER PER ME466-O003, DATED 10 NOVFMBER 1970.

2.4 VOLTAGE WAVE FORM. THE LINE-TO-LINE VOLTAGE WAVE FORM OF THE QUASI- SQUARE WAVE SHALL BE PER FIGURE 1. "SPIKE"l DISTORTIONS SHALL NOT EXCEED 40 VOLTS PEAK RELATIVE TO NOMINAL WAVE FORM. THE LINE-TO-LINE VOLTAGE WAVE FORM OF THE FILTERED OUTPUTS SHALL BE SINUSOIDAL WITH 10 PERCENT MAXIMUM DISTORTION.

2.5. REGULATION. THE OUTPUT OF THE INVERTER SHALL BE CURRENT/VOLTAGE REGULATED AS FOLLOWS. THE IMPEDANCE OF THE GYRO MOTORS SIGNIFICANTLY INCREASES WHEN THEY REACH SYNCHRONOUS SPEED. THE INVERTER SHALL SENSE THIS AND REACT AS A FUNCTION OF THIS LOAD CHANGE BY CHANGING FROM AN OUTPUT CURRENT R3EGULATED MODE (DURING GYRO START AND RUNUP) TO AN OUTPUT VOLTAGE REGULATED MODE (AT SYNCHRONOUS SPEED) WITHIN THE FOLLOWING LIMITS:

(A) CURRENT REGULATED MODE: .850 RMS AMPS PER PHASE MINIMUM; THE MAXIMUM CURRENT SHALL BE LIMITED AS NECESSARY TO MEET THE REQUIREMENTS OF PARAGRAPH 2.5 (THE CURRENT WAVE FORM APPROXI- MATES A SINE WAVE DUE TO THE INDUCTIVE LOAD).

(B) VOLTAGE REGULATED MODE: 275 VOLTS PEAK-TO-PEAK, LINE-TO-LINE MINIMUM; MAXIMUM VOLTAGE SHALL BE LIMITED AS NECESSARY TO MEET THE REQUIREMENTS OF PARAGRAPH 2.5.

(C) REGULATION (CURRENT AND/OR VOLTAGE) SHALL BE BASED ON THE AVERAGE OF THE THREE OUTPUT PHASES. WITH A BALANCED LOAD, INDIVIDUAL LINE-TO-LINE VOLTAGES AND/OR LINE CURRENTS SHALL NOT DEVIATE MORE THAN 5 PERCENT FROM THEIR RESPECTIVE AVERAGES.

2.6 EFFICIENCY. THE REQUIRED DC INPUT CURRENT FOR VARIOUS INPUT VOLTAGES SHALL NOT EXCEED THE FOLLOWING VALUES:

(A) IN THE CURRENT REGULATED MODE (FROM GYRO MOTOR START THROUGH GYRO MOTOR RUNUP):

AT 22 VDC INPUT ...... 5.0 AMPS

AT 28 VDC INPUT...... 4.6 AMPS 5-

TOLERANCES. EXCEPTAS NOTED TOLERANCES ON .3 11U :0 : :0.0 TU .7. +. -. W l D041THRU, 13D +.02 -.081 .751 TMU 1.000,+.007 -. 31 CODE WENT NO. ANGLES *V' DECI MAl$S - '03 MOtES NO1D 131THI*.22 ,+.= -. 31 1.001 ThU 2.0W 4.010 -. 01 03953 • .00 0R11L .230 THU ..0 :4.04 -.001

Dg BY J. WOLTZ -17-7( SPACE DIVISION CON7U- PART NORTfH Al4 CAN ROCKWELL CORPORATION CH(BNY i F 4t *YM_95-OO11 AINIBlTIR, THREE-PHASE

SPEC CONTROL DRAWING slMET 3 of 6

E-3 Space Division 0 NorthAmencan Rockwell

L-. - ONE CYCLE

t3

t = t3 = 2t2 2t4

FIGURE 1 - LINE-TO-LINE VOLTAGE WAVE FORM

TOLERANCES, EXCEPTASNOTED TOLERANCES ON .013 TH" .040 +.I -. 0I .501 THRUT W.+.005 -.S 041 TIHU. 130: +.M2 -. 011 .732THtU i., +.007 -. OD CODECENT NO. CI. Xx ­*.010 'DLL'NOE .131 THRU .22 0 -. 001 1.001 THtU 2.WD0. +.DI -.001 03M5 ANLS00ANGLES Mr DECMAISE .,YXX -*.03.P HOLESDIL NOTED .230* THRU .MI:n.ms.0J04 -. 001 .1 1t2~.I-w 35 I*y J. WOLTZ -17-7 SPACE DIVIION ITR0L PART NORTH AMERICAN ROCKWELL CORPORATIu P CHWy14 L fo SCIAIM• Y. CMSAA 04 AfIOID BY ME495-O0 nVERTER, THREE-PHASE SPEC CONTROL DRAWING SherEE 4 OF 6

E-4 11dil Space Division @T NorthAmencan Rockwell

AT 34 VDC INPUT ...... 4.1 AMPS

(B) IN THE VOLTAGE REGULATED MODE (WITH GYRO MOTORS AT SYNCHRONOUS sPEED):

AT 22 VDC-INPUT ...... 2.8 AMPS

AT 28 VDC INPUT ...... 2.4 AMPS

AT 34 VDC INPUT ...... 2.1 AMPS

WITH THE GYROS AT SYNCHRONOUS SPEED, THE INPUT CURRENT OSCILLATES (GYRO MOTORS "HUNTING"); THE ABOVE MAXIMUM VALUES ARE THE AVERAGES OF THE OSCILLATORY MAGNITUDES.

2.7 FLIGHT LIFE. THE FLIGHT LIFE OF THE INVERTER, INCLUDING DESIGN MARGINS, SHALL OF ONE YEAR NON-OPERATING, AND 120 HOURS OF OPERATION.

3. ENVIRONMENTS

3.1 GENERAL. THE APPLICABLE OPERATING/NON-OPERATING GROUND AND FLIGHT ENVIRONMENTAL REQUIRE4ENTS OF SPECIFICATION MC823-O007 SHALL APPLY, AS AMENDED BY THE FOLLOWING: (A) THE INVERTER SHALL BE CONSIDERED AS A CLASS II ITEM INSOFAR AS THE NATURAL FLIGHT ENVIRONMENTS ARE CONCERNED, (B) THE NON-OPERATING FLIGHT TEMPERATURE SHALL BE 30 TO 110 DEGREES F, AND (C) THE OPERATING FLIGHT TEMPERATURES SHALL BE 30 TO 90 DEGREES F.

4. IDENTIFICATION AND MARKING

4.1 IDENTIFICATION OF PRODUCT. IDENTIFICATION AND MARKING OF THE INVERTER SHALL BE AS SPECIFIED IN SPECIFICATION MC823-0007. 5. QUALITY ASSURANCE PROVISIONS

5.1 TESTING REQUIREMENTS. THE QUALITY ASSURANCE PROGRAM (ACCEPTANCE AND/OR QUALIFICATION TESTING) NECESSARY FOR ASSURANCE THAT THE INVERTER QUALIFIES FOR THIS APPLICATION SHALL BE DETERMINED BY THE SUBCONTRACTOR UTILIZING THE CRITERIA SPECIFIED IN SECTION 4 OF SPECIFICATION MC823-0007. PRIOR TO IMPLEMENTATION, THE TEST PROGRAM SHALL BE SURMITTED TO NR/SD FOR APPROVAL.

5.1.1 INDIVIDUAL PERFORMANCE TESTS. THE FOLLOWING INDIVIDUAL PERFORMANCE TESTS AND INSPECTIONS SHALL BE INCLUDED AS PART OF THE QUALITY ASSURANCE PROGRAM.

(A) EXAMINATION OF PRODUCT

(B) INSULATION RESISTANCE TEST

EXEP D TOLEXANCES ON .013 THU .0,: .1 001.. .01THRU 710 . .. T T041 THRU.130: +.2 -. 0I .7S1THRJI.( DO+. 0*-.U CODEDENTNO. AxES DCALS .XX*.l HOLS NOTED .:.-I 1.I THOU2.-..0X1-.001 03 .23lIEU.919: +.0S4 -61 D BY J. WOLTZ INORtTH -17-7( AMEICANm1ACE ROCKWELL DIVISION CORtPORATION CONTR PART

CHJ Y VE BYV L AFO M I N E 49 5 -Mf- INVERTER, THRE-PHASE SPEC CONTROL DRAWING SHEET '5 of 6

E-5 0 i Space Division North Amencan Rockwell

(C) PERFORMANCE VERIFICATION TESTS RELATED TO FREQUENCY, OUTPUT WAVE-FORM, PHASE ROTATION, REGULATION, AND EFFICIENCY. THESE TESTS SHALL BE CONDUCTED WITH USE OF A SUPPLIER-FURNISHED STATIC LOAD BANK WHICH SIMULATES THE ACTUAL COMBINED GYRO MOTOR AND RESOLVER LOAD AT THREE CONDITIONS: (A) AT GYRO MOTOR STARTING, (B) AT MAXIMUM INPUT POWER CONDITION DURING MOTOR RUNUP, AND (C) AT SYNCHRONOUS SPEED OF THE GYRO MOTORS. SUITABILITY OF THIS LOAD BANK AS AN EQUIVALENT SIMULATED LOAD SHALL BE APPROVED BY NR/SD.

6. PREPARATION FOR DELIVERY

6.1 PACKAGING, PACKING, AND MARKING. PACKAGING, PACKING, AND MARKING OF THE INVERTER SHALL BE AS SPECIFIED IN SECTION 5 OF SPECIFICATION MC823-0007.

TOLERANCES, EXCEPT AS NOTED TOLERANCES ON .013 THtU .040: ,.* -. I .001ITHU 7W.+.O.D

A* A.0GLS.0'CWAL$ 3-*'0 HOLESNOTED .041TMRU.|3IO .I-.0 .751TIIJ,. .. 7-.0I COCE I0TIO. A S.+ 0D'iLL .131TRU..2:x.00-.01 1.001DL*tU2.0:I.'+.0100 03953 i at J. WOLTZ fl-17-7 SPACE DIVON CONTROL PART CHKIY NORTH2214 LJ(LAMERICAN I OAXAO ROCKWELL* OD*MY, CMLIkACORWRATION N1M)4 B APROV INVERTER, THREE-PHASE ME495-0f

SPEC CONTROL DRAWING ET 6 of 6

E-6 Albk Space Division 9I NorthAmencan Rockwell

PREPARED BY CODE IDENT. NO.. 03953 NUMBER ;MC823-0007 j. Woltz/E. Bondurant SPACE DIVISION TYPE GENERAL APPROVALS NORTH AMERIC4N ROCKWELL CORPORATION PROCIJRMENT 12214 LAKEWOOD BOULEVARD " DOWNEYCALIFORNIA 90241 DATE 6 'c/_ . e t-a-7~5 NOVEMBE 1970 SPECIFICATION SUPERSEDES SPEC. DATED: REV. LTR. "PAGE I of 21

TITLE ENVIRONMENTAL AND GENERAL DESIGN REQUIREMENTS -OR REMOTE MANEUVERING UNIT SUBSYSTEMS AND OMPONENTS, GENERAL SPECIFICATION FOR

FORM M 131-H-1 REV 11-67 (0- INDICATES CHANGE)

E-7 Space Division 9D North American Rockwell

SPACE DIVISION NORT1II \MERICAN; ROCKWELL CORPOR MON 12214 LAKEWOOD BOULEVARD - DOWNEY, CALIFORNIA 90241 CODE-IDENT. NO. 03953

MC823-0007 PAEi

TABLE OF CONTENTS

1. SCOPE 2

1.1 Scope 2

2. APPLICABLE DOCUMENTS 2

2.1 Applicability 2

3. REQUIREMENTS 3

3.1 Materials, Parts, and Processes 3 3.1.1 Materials and Parts 3 3.1.2 Dissimilar Metals 3 3.1.3 Fungus Nutrient Materials 3 3.1.4 Extraneous Materials 3 3.1.5 Hazardous Materials 4 3.1.6 Hazard Prevention 4 3.1.7 Cleanliness 4 3.2 Design 4 3.2.1 Weight 4 3.2.2 Size 4 3.2.3 Maintainability 4 3.2.4 Structural Requirements 4 3.2.5 Shelf Life 4 3.2.6 Age Control 4 3.2.7 Lubrication 4 3.2.8 Electrical Design 4 3.2.8.1 General Specifications 4 3.2.8.2 Electrical Grounding 5 3.2.8.3 Conductivity 5 3.2.8.4 Insulation Resistance 5 3.2.8.5 Electromagnetic Interference (EM) Compatibility 5 3.3 Construction 5 3.3.1 Threads and Fasteners 5 3.3.2 Electrical Bonding 5 3.3.3 Soldering 5 3.3.4 Crimping 5 3.3.5 Retainer Rings 5 3.3.6 Finish 5 3.3.7 Identification and Marking 5 3.3.8 Workmanship 6

FORM M 131-H-2 REV 11-67

E-8 'Ib Space Division North American Rockwell

SPACE DIVISION NORTH AMERIC \N ROCKWELL CoRFOR MON 12214 LAKEWOOD BOULEVARD * DOWNEY, CALIFORNIA 90241 CODE ]DENT. NO. 03953

1NUMBER REVISION7LETTERI MC823I0007 PAGE ii

TABLE OF CONTENTS (Cont'd)

PARA. TITLE PAGE

3.4 Environments 6 3.4.1 Natural Environments 6 3.4.1.1 Ground Environments (Natural) 6 3.4.1.2 RMU Flight Environments (Natural) 7 3.4.2 Induced Environments 9 3.4.2.1 Ground Environments (Induced) 9 3.4.2.1.1 Non-operating Environments 9 3.4.2.1.2 Operating Environments l1 3.4.2.2 Flight Environments (Induced) 1l 3.5 Reliability 13 3.5.1 Useful Life 13 3.5.2 Flight Life 13 3.5.3 Reliability Verification 13

4. QUALITY ASSURANCE PROVISIONS 14

4.1 Quality Assurance Program 14 4.2 Test Program 14 4.3 Traceability 14 4.4 Classification of Tests 14 4 .5 Definition of Test Classifications 14 4.5.1 Acceptance Tests 14 4.5.2 Qualification Tests 15 4.6 Test Conditions 15 4.7 Test Levels 15 4.8 Measurement Tolerances 15 4.9 Test Description and Sequence 16 4.9.1 Examination of Product 16 4.9.2 Performance Test 16 4.9.3 Leak Detection 16 4.9.4 Structural Dynamic Tests 16 4.9.4.1 Sinusoidal Vibration 16 4.9.4.2 Random Vibration 17 4.9.4.3 Acoustic Noise 17 4.9.5 Acceleration - Steady State 18 4.9.6 Thermal Vacuum 18 4.9.6.1 Corona Test 18 4.9.6.2 High Temperature Soak 18 4.9.6.3 Low Temperature Soak 19 4.9.7 Electromagnetic Field Measurement 19

FORM M 131-H-2 REV t1-67

E-9 Space Division North American Rockwell

SPA.CE DIVISION ORIIl 1EI.i,\MERI( \N ROCKI CORPORATION 12214 LAKEWOOD BOULEVARD - DOWNEY. CALIFORNIA 90241 CODE IDENT. NO. 03953

NUMBER H83007REVISION LETTER PG

TABLE OF CONTENTS (Cont'd)

PARA. TITLE PAGE

5. PREPARATION FOR DELIVERY 20

5.1 General Packaging Requirements 20 5.1.1 Preservation and Packaging 20 5.1.2 Packing 20 5.2 Packaging Design Criteria 20 5.2.1 Size and Weight 20 5.2.2 Drop Height Criteria 20 5.3 Marking for Shipment 20

6. NOTES 21 6.1 Definitions 21 6.1.1 NR/SD 21 6.1.2 Subcontractor 21

FORM M 131-H2 REV tI-67

E-10 @1 Space Division NorthAmerican Rockwell

SPACE DIVISION N'

1. SCOPE

1.1 Scope. This specification establishes the general design and environmental requirements for the Remote Maneuvering Unit (EMU) subsystem and components. Thi' document shall be used for reference and/or applicability in all RMU subsystem and components procurement documents (Specification Control Drairings (SCD)).

2. APPLICABLE DOCUMENTS

2.1 Applicability. The following documents, of the issue specified, or if undated the latest issue in effect, shall form a part of this specification to the extent speci­ fied herein. Wlhen the requirements of this specification and the documents referenced herein are in conflict, the requirements of this specification shall be considered a superseding requirement:

SPECIFICATIONS

Military

MIL-B-5087A (1) Bonding Electrical, Lighting 29 January 1958 Protection for Aerospace Systems

MIL-E-54OOK Electronic Equipment, Aircraft, 7 December 1966 General Specification for

MIL-S-7742B Screw Threads, Standard, Optimum 2 February 1968 Selection Series, General Specifi­ cation for

MIL-E-8189B (ASG), (2) Electronic Equipment, Guided Missiles, 1 July 1964 General Specification for

MIL-S-8879A Controlled-Radius Root with Increased 8 December 1965 Minor Diameter, General Specification for

National Aeronautics and Space Administration

MSC-NPC 200-4 (ASP0-S-6A) Quality Requirements for Hand August 1964 Soldering of Electrical Connections

MSC-SPEC-Q-1 Crimping of Electrical Connections, Specification for

North American Rockwell Corp., Space Division NR/SD

ST02OMA00O7A Identification and Traceability for 5 November 1970 Subcontractors Remote Maneuvering Unit (RMU) Equipment STO802GTOO01B Supplier Quality Program 15 October 1968

FORM M 131-H-2 REV 11-6

E-11 l. S p ~c e :D iv is lOn NorthAmredcan Rockwell

SPACE DIVISION NORTH IAMERIC \N ROCKNWElI. CORPOR \TION 12214 LAKEWOODBOULEVARD • DOWNEY, CALIFORNIA 90241 CODE IDENT. mO. .03953 NUBER C823-0007AE 3

STANDARDS

Military

MIL-STD-130B (1) Identification Marking for 7 February 1964 U.S. Military Property

MIL-STD-129D (11) Marking for Shipment and Storaga 4 April 1969

MIL-STD-794B Parts and Equipment, Procedure 17 March 1969 for Packaging and Packing of

MIL-STD-461 Electromagnetic Interference 31 July 1967 Characteristics Requirements for Equipment

MIL-STD-Dwg. M333586A Metals, Definition of Dissimilar' (ASG) 16 December 1958

3. REQUIREMENTS

3.1 Materials, Parts, and Processes.

3.1.1 Materials and Parts. Materials and parts qualified for previous spacecraft missions shall be used wherever possible. Whenever use of such materials and parts is not possible, standard parts, Air Force-Navy (AN), Military Standard (MS), ci joint Air-Force-Navy (JAN) materials and parts shall be used wherever they are -suitable for teb1~service, and provided they conform to the requirements of this specification. Materials and parts used shall withstand the environment to which they will be exposed

>wtthout deterioration. ­

3.1.2 Dissimilar Metals. Dissimilar metals as defined in Standard-MS 33686 (ASG) shall not be used in contact, unless suitably protected against electrolytic corrosion. The protection used shall offer a low impedance path to radio frequency currents.

3.1.3 Fungus Nutrient Materials. The use of fungus nutrient materials shall be avoided. When such materials are used,, they shall be adequately treated or hermetically sealed with a suitable fungicidal agent to prohibit fungus growth.

3.1.4 Extraneous Materials. Materials shall not emit particles or substances generated within themselves as a result of wear. Surfaces shall be sufficiently smooth and wear resistant to minimize the generation of metal-to-metal and other wear particles. All finishes shall be compatible with the operation mediums,and 'shall be capable of with­ standing exposure to the environmental conditions referenced herein for the life of the respective unit. Zinc and cadium plating shall not be used.

WORM M 131-H-Z REV lI-67

E-12 AI k Space Division @L North Amencan Rockwell

SPAC.E DIVISION NOR IIII XMERIC \N ROCKWEIl., CORPOR MON 12214 LAKEWOOD BOULEVARD • DOWNEY, CALIFORNIA 90Z41 CODE IDENT. NO. 03953 NUMBER i REV°Sr LETTER007l PAGE4

3.1.5 Hazardous Materials. The use of hazardous materials (e.g., flammable, explosive, toxic, etc.) shall be avoided. Where use of such materials is considered necessary, appropriate drawyings or documents describing the material's hazardous characteristics and criteria shall be submitted to NR/SD for approval prior to design implementation.

3.1.6 Hazard Prevention. Subsystems and components shall be designed to prevent hazards of electrical shock, fire, explosion, toxicity, etc., to personnel, equipment, and facilities during assembly, test, and operations.

3.1.7 Cleanliness. The RMU equipment shall be cleaned to remove all foreign materials such as dirt, dust, grease, lint, or any other matter that will degrade its performance.

3.2 Design.

3.2.1 Weight. Weight shall be maintained to an absolute minimum consistent with good design. The weight shall not exceed the weight specified in the applicable SCD.

3.2.2 Size. The size of the RMU equipment shall be maintained to a minimum consistent with good design. The size of the unit shall not exceed the envelope dimensions specified in the applicable SCD.

3.2.3 Maintainability. Mechanical, functional, and electrical interchangeability shall exist between replaceable units, regardless of the supplier. Interchangeability shall require that a substitution of the units be affected without physical or electrical modifications to any part of the unit. Requirements for ease of maintenance shall be limited to provide specific interchangeability (high density packaging requirements shall take precedence over requirements for ease of maintenance).

3.2.4 Structural Requirements. A structural safety factor of 1.50 shall be used as the minimum value in structural design to determine ultimate loads. Ultimate loads shall be limit loads multiplied by the safety factor.

3.2.5 Shelf Life. The RMU equipment shall be designed to have a minimum shelf life of 3 years under normal warehouse conditions.

3.2.6 Age Control. Elastomers shall not be used that require age control.

3.2.7 Lubrication. The REM equipment shall noL require lubrication for the life of the equipment.

3.2.8 Electrical resign.

3.2.8.1 General Specifications. Electrical equipment shall be compatible with applicable requirements of Specifications MIL-E-5400 and MIL-E-8189, with the latter taking precedence.

121-H-2 I13-S EV

E-13 Space Division 91 North American Rockwell

SPACE DIVISION NORTII \MERIC\N ROCKWE.L :ORPOR \TION 12214 LAKEWOOD SOULEVARD • DOWNEY, CALIFORNIA 90241 CODE ]DENT. HO 03953 NUMBER REVISION LETTER P c82 I I I07IPAE

3.2.8.2 Electrical Grounding. All external metallic parts of the equipment shall be at ground potential. All electrical returns shall be isolated from the equipment case.

3.2.8.3 Conductivity. Materials used in electronic or electrical connections shall have such characteristics that during specified environmental conditions there shall be no adverse effect upon the conductivity of the connections.

3.2.8.4 Insulation Resistance, The insulation resistance from any electrical interface of the equipment to the case shall exceed 100 megohms of 100 vdc, unless otherwise specified in the equipment SCD.

3.2.8.5 Electromagnetic Interference (EI)Compatibility. Operation of the equipment and/or subsystem shall not interfere with the functioning of the electronic equipment of the spacecraft due to spurious electromagnetic radiation. The requirements of Standard MIL-STD-461, Class IIA equipment shall apply.

3.3 Construction.

3.3.1 Threads and Fasteners. Screir thread shall be in accordance writh Specifica­ tions NIL-S-7742 or NIL-S-8d79 for fastener ultimate tensile strengths below 160,000 pounds per square inch (psi). Specification MIL-S-8879 shall be used for fastener -ultimate tensile strengths 160,000 psi and above. External threads in accordance with Specification MIL-S-8879 for fastener tensile strength 160,000 psi and above shall be produced by a single thread rolling process after final heat treatment. Removable, externally threaded bushing type, corrosion-resistant steel inserts shall be used in tapped holes in aluminum alloys and magnesium alloys where frequent removal of the threaded part wrill be encountered in normal service or maintenance. Coiled wire type inserts shall not be used.

3.3.2 Electrical Bonding. Where electrical bonding is used the bonding shall be in accordance with Specification MUL-B-508.

3.3.3 Soldering. Solder, preparation for soldering, and soldering shall be in accordance with NPC 200-4 as supplemented by ASTO-S-6A.

3.3.4 Crimping. Crimping of electrical connections shall be in accordance with the applicable portions of Specification MSC-SPEC-Q-1.

3.3.5 Retainer Rings. No retainer rings or snap rings shall be used in the construction of RMU equipment.

3.3.6 Finish. Finishing of RMU ecuipment shall be compatible with the specified performance, and shall withstand the environments to wrhich it will be exposed without deterioration.

3.3.7 Identification and Iarking. The RMU equipment shall be durably and legibly marked (or tagged if sufficient area is not available for marking) in accordance writh Standard MIL-STD-130, and shall include the folloiring information:

FORM M 131-MA-2 REV 11-67

E-14 A b Space Division 9L North Amencan Rockwell

SPACE DIVISION NORTI INIERI(AN ROCK1%EILL CORPORATION 12214 LAKEWOODBOULEVARD • DOWNEY,CALIFORNIA 90241 CODE IDENT. NO- 03953 [NUMBER MC82LTE7 R PAGE 6

Item Name

NR/SD Control No. (Allow for 15 digits)

NH/SD Inspection Serial No.

Manufacturer

Manufacturer Part No.

Manufacturer Serial No.

Date of Manufacture

The subcontractor shall insert the item name as shown in the title of the equipment SCD, manufacturer's name, manufacturer's serial and part number, date of manufacture, and the US as shown. When required, NR/SD personnel will apply the inspection serial number. This will be accomplished either at the subcontractor's facility or after arrival at NR/SD.

3.3.8 Workmanship. The RMJ equipment shall be constructed and finished giving particular attention to marking, neatness and thoroughness of assembly, welding, solderin plating, painting, machining, and fit and adherence to specified dimensions and tolerances. 3.4 Environment s.

3.4.1 Natural-Environments.

3.4.1.1 Ground Environments (Natural). The RMU spacecraft equipment will not be operated on the ground within natural environmental conditions (all ground operating modes irill occur in conditione4 induced environments). Accordingly, the following represent the natural environmental extremes which may be encountered by the RMU and EMU equipment in a nonoperating condition during transportation, ground handling, and storage. EMU equipment may be protected by suitable packaging for transportation and storage if these environments exceed the equipment flight design operation require­ ments. The equipment shall be capable of meeting the operating requirements of the applicable SCD after exposure, while protected by its normal packaging, to these environments:

(a) Temperature (Air)

Air Transportation -45 to +140 F for 8 Hours

Ground Transportation -20 to +145 F for 2 Weeks

Storage +25 to +105 F for 3 Years

FORM M 13i-H-2 REV 11-67

E-15 Q Space Division North Amencan Rockwell

SPACE DIVISION NOR II MNISRIC \N ROCKWElL CORPOR \ IO\ 12214 LAKEWOOD BOULEVARD * DOWNEY,CALIFORNIA 90241 CODE IDENT. NO. 03953 NUMBER HC2-00 LETTER 1 MC823-0007 PG

(b) Pressure

Air Transportation Minimum of 3.47 psia for 8 Hours (35,000 ft. altitude)

Ground Transportation Minimum of 11.78 psia for 3 Years (6,000 ft. altitude)

(c) Humidity 0 to 100 percent relative Humidity, including conditions wherein condensation takes place in the form of water or frost for at least 30 days

(d) Sunshine Solar radiation of 360 BTU per square foot per hour for 6 hours per day for 2 weeks

(e) Rain Up to 0.6 inch per hour for 12 hours. 2.5 inch per hour for 1 hour

(f) Sand and Dust As encountered in desert and clean beach areas. Equivalent to 140-mesh silica flour with particle velocity up to 500 feet per minute and a particle density of 0.25 grams per cubic foot.

(g) Fungus As experienced in Florida climate. Materials will not be used which will support or be damaged by fungi.

(h) Salt Spray Salt atmosphere as encountered in coastal areas, the effect of which is simulated by exposure to a 5 percent salt solution by weight for 48 hours.

(i) Ozone 3 years exposure, including 72 hours at 0.5 PPM, 3 months at 0.25 PPM, and the remainder of 0.05 PPM concentration.

3.h.1.2 RMU Flight Environments (Natural). The following represent the natural environmental extremes for the EMU spacecraft equipment as experienced during the various flight mission phases. Unless otherwise specified in the specific SCD, these environments apply to both operating and non-operating conditions:

Equipment mounted on the interior of the spacecraft will be partially protected from some environments. Such environments are identified in the following as applicable to Class I equipment only; environments without such a designation are applicable to both Class I and Class IT equipment. The applicable SCD identifies each individual procured item as Class I or Class II equipment. The RMU equipment shall be capable of meeting the operating requirements of the applicable SCD during and after exposure to the following environments:

FORM M 131-H-2 REV 11-67

E-16 Al Space Division 'AL North Amencan Rockwell

SPACE DIVISION NOR Il I AMERIC \N ROCK EI.. CORPOR VI(N 12214 LAKEWOOD BOULEVARO . DOWNEY.CALIFORNIA 90241 MC2-0O_oo CODE IDENT. NO. 03953 NUMBER j REVISIONEIo° LETRPAGE8E;I I I O

(a) Pressure During launch the reference earth atmosphere at Cape Kennedy, Florida, shall be in accordance with NASA TM X-53139. During subsequent mission phases the environmental pressure will be less than i0- 1 3 MM Hg. (b) Electromagnetic The sources of electromagnetic radiation Radiation. These presented below impinge on the exterior environmental criteria surfaces of the RMU in logical combination. are applicable to com­ ponents identified as Class I only. 2 Solar Flux (All Wave Length) 442 BTU/FT -Hour

Earth Emission 73 BTU/FT 2 -Hour (Excluding Reflection)

Earth Average Albedo Approximated by that from a 6000 K (Visual Range) Black Body Earth Average Albedo 0.39 (Total Solar Spectrum)

Space Sink Temperature Zero Degrees Rankine

(c) Meteoroid. The flux-mass sporadic meteoroid model This environment is appli- for RMU exterior exposure shall be in cable to components identi- accordance with NASA TM X-53521. tied as Class I only.

(d) Nuclear Radiation

(i) Trapped Radiation - Radiation levels in the Van Allen and artificial belts will use proton and electron fluxes obtained with the Goddard Orbital Flux Code.

(2) Galactic Cosmic Rays - Galactic Cosmic Ray doses range from 6 to 20 Rads/Yr.

FORM M 131-H-2 REV 1t-67

E-17 Alk Space Division 4 North Arnencan Rockwell

SPACE DIVISION NO\KTl II \N ROCKW.I. (tRIOR \ I 12214 LAKEWOODBOULEVARO • DOWNEY,CALIFORNIA 90241

- CODE ]DENT. NO. 03953 N*UMBER REVISION LETTER

(3) Solar Particle Events - The Solar Particle environments in terrestrial space, within the Earth's magnetic field, apply for rigidities above the cut-off rigidities given by: (1 )) N 2.49 29x109x 10 2 * COS3A0%.21-3 - 2 + COS33

where N = Particle's cut-off rigidity, BV

h = Altitude, KM

= Geomagnetic Latitude

Solar particle events will be considered to contain solar produced alphas and protons irith equal rigidity spectra.

3.4.2 Induced Environments.

3.4.2.1 Ground Environments (Induced). Two types of ground environments must be considered: (a) conditions which apply irith the RMU equipment operating, and (b) conditions which apply when the RMU equipment is not operating.

3.4.2.1.1 Non-operating Environments. The following represent the induced environ­ mental extremes which may be encountered by RMU equipment in a non-operating condition during transportation, ground handling, and storage. RMU equipment may be protected by suitable packaging for transportation and storage if these environments exceed the equipment flight design operation requirements. The equipment shall be capable of meeting the operating requirements of the applicable SCD after exposure, while protected by its normal packaging, to these environments:

(a) Shock

(1) Packaging - The design shock levels applicable to protective packaging, or directly to equipment when packaging is not used, shall be those resulting from free fall drops as presented in the table below. The protective packaging, if used, shall attenuate shock such that packaged RMU equipment will not be exposed to shock levels exceeding the equipment flight design operating requirements.

FORM M t31-H- REV 1]-67

E-18 Qdl Space Division North American Rockwell

SP WOEDIVISION NORFI'I \MERIC \N ROCKWELL CORPOR \TION 12214 LAKEWOOD BOULEVARD * DOWNEY, CALIFORNIA 90241

CODE [BENT mO. 03953

NUMBER 1 MC 2 -0 7REVISiONLETTER PAG 10 1

DIMENSIONS** DROP HEIGHT WEIGHT* NOT EXCEEDING FREE FALL EDGEWISE CORNERWISE

Less than 50 lbs. 36 Inches 30 Inches

50 to 100 lbs. 48 Inches 21 Inches

100 to 150 lbs. 60 Inches 18 Inches

150 to 200 lbs. 60 Inches 16 Inches

200 to 600 lbs. 72 Inches 36 Inches 36 Inches

Over 600 lbs. Over 72 In. 24 Inches 24 Inches

* Weight of equipment and package or containers (if used) ** Dimensions along any edge or diameter

(2) Equipment - The design shock levels applicable directly to Rb4U equipment, hich utilizes protective packaging, for which no other design operating requirements are established, shall be as presented in the following table:

WEIGHT SHOCK htrEL - TIME (POUNDS)* (8) -(MILLISECONDS)

Less than 250 30 11 + 1 (half-sine waveform)

250 to 500 24 11 + 1 (half-sine waveform)

500 to 1,000 21 11 + 1 (half-sine waveform)

Over 1,000 18 11 + 1 (half-sine waveform)

* Weight of equipment and package or containers (if used) ** As experienced in any direction

(b) Vibration - Sinusoidal as experienced in any direction.

WEIGHT 5 to 26.5 26.5 to 52 52 - 500 (POUNDS)* CPS CPS CPS (Inch DA)

Less than 50 + 1.56 g 0.043 + 6.0 g 50 to 300 + 1.30 g 0.036 + 5.0 g 300 to 1,000 + 1.30 g 0.036 NA Over 1,000 + 1.04 g 0.029 NA

Weight of equipment and package or container (if used)

FORM M 131-H-2 REV 11-67

E-19 Space Division @0 North American Rockwell

SPACE DIVISION NORTH \fMERIC N ROCKW'EI.I. CORPORATION - 12214 LAKEWOOD BOULEVARO- DOWNEY,CALIFORNIA 90241' CODE IDENT.NO. 03953 NtJIBRI--PAGE ll MC823--0007TAEl NUMBER M02-07REVISION LETTER

3.4.2.1.2 Operating Environments. Operation of the BMU equipment (other than in environmental testing) will occur in sheltered areas only, such as inside of buildings or inside of protective canopies/shrouds on the launch pads atop the booster. In all cases the environments vill be controlled in such a manner that they will at no time be more severe than the corresponding environments occurring during flight phases. Within the canopy/shroud envelope, wrhen the RMU spacecraft is mated to the launch vehicle, the environmental conditions are controlled to provide the following environmental conditions:

(a) Pressure: 28 - 32 In. Hg.

(b) Temperature: 77 + 50 F

(c) Humidity: Less than 50 percent

"(d) Filtration: Filters up to 99.97 percent of all particles in excess of 0.3 microns

Within buildings (vehicle test and assembly areas) the operating environmental conditions are controlled to provide the following environmental conditions:

(a) Pressure: 28 - 32 In. Hg.

(b) Temperature: 72 + 100 F

(c) Humidity: Less than 80 percent

(d) Filtration: Less than 1 percent of all particles are in excess of 0.5 microns

3.4.2.2 Flight Environments (Induced). The following represents the induced environments for the EMU equipment as experienced during various flight mission phases. Unless otherwise specified in the specific SCD, these environments apply to both operating and non-operating conditions, The equipment shall be capable of meeting the operating requirements of the applicable SCD during and after exposure to these environments:

(a) Acoustics

The acoustic environment is comprised of broad band random pressure fluctua­ tions resulting from booster engine noise and boundary level turbulence. The frequency distribution of noise at various maxima (lift-off, transonic and maximum dynamic pressure) are as tabulated below:

FORM .14I1-i--2 RE ,-6

E-20 A& Space Division 9 NorthA6nencan Rockwell

SP\CFE DIVISION NOR Ill \\IERXAN ROCKWEL. (ORPOR0 10% 12214 LAKEWOODBOULEVARD • DOWNEY,CALIFORNIA 90241 CODE IDENT.NO. 03953 UMER mc823-0007 REV° LETTER IPAE12 I

Octave Band Center Sound Pressure Level 2 Frequency Hz db (Ref., 0.0002 DTnes/CM ).

15.8 123

31.5 123

63 125

125 130

250 133 500 134

1000 13k

2000 131

4000 127

8000 121

Overall 140

(b) Acceleration

The maximum steay state acceleration occurs during the firing of the spacecraft Apogee motor. The spacecraft is spinning at a nominal rate of 60 RPM at this time. The actual acceleration level is, therefore, a resultant of the axial acceleration of 13.97g and the radial acceleration. This radial acceleration is a function of the spin rate (60 RPM) and the radial distances of the component from the spin axis. The resultant acceleration may be computed as follows where r is the radial distance in inches and N = spin rate in RPM = 60.

5 2 gr 13972 + E2.84 X 10 ) X r N

When no radial distance is specified in the SCD, 25.5 inches shall be used.

(c) Vibration

(1) Thrust Axis - Sinusoidal Vibration Frequency (Hz) Accleration (0 to Peak) 5 - 11 0.32 inch double amplitude 11- 17 +2.Og 17 - 23 + 4.o g 23 - 200 + 1.5 g 200 - 2000 + 3.3 g

FORM M I31-H-2 REV 11-67

E-21 0 Space Division North American Rockwell

SPACE DIVISION NOR"I II \\IERIC %NROCKWELL CORPOR TION 12214 LAKEWOODBOULEVARD - DOWNEY,CALIFORNIA 90241 CODE IDENT. NO. 03953 NUMBER M0823-0007 REVISION LETTER PAGE 13

(2) lateral Axis - Sinusoidal Vibration

Frequency (Hz) Acceleration (0 to Peak)

5 - 7.5 0.52 inch double amplitude 7.5 - 13 + 1.5 g

13 - 200 + 1.O g 200 - 2000 + 3.3 g

(3) Thrust Axis & lateral Axis - Random Vibration Acceleration 2 ) Density (g /Hz) (g-R Frequency (Hz) Power Spectral 20 - 300 .0013 to 0.02 Increasing from 6.1

20 HS at a rate of 3 db/octave

300 - 2000 0.02 6.1

(d) Temperature

The temperature extremes shall be specified by the specific SCD for both operating and non-operating conditions.

(e) Electromagnetic Interference.

The susceptibility and compatibility requirements of the equipment shall be as specified in the specific SCE), where applicable.

3.5 Reliability. Particular emphasis shall be placed on elimination of single failure points affecting primary operating modes. The equipment shall be designed to meet the performance requirements while exposed to the environments specified herein for the durations specified in the following paragraphs.

3.5.1 Useful Life. The period from time of assembly of the equipment until its designed operational application has been completed shall be 3 years including storage and all acceptance, qualification, and checkout test operations. The equipment will be exposed to component, subsystem and spacecraft acceptance, qualification and checkout tests prior to its use as a flight article. Therefore, the design shall be compatible with the required flight life after exposure to all these test conditions.

3.5.2 Flight Life. The total failure free period of time (including design margins) that the equipment shall function during flight shall be as specified in the specification control drawing for non-operating and operating conditions.

3.5.3 Reliability Verification. Reliability will not be demonstrated by testing or analytical means unless otherwise specified by the purchase order. IM/SD may, at their discretion, review the detailed designs to ascertain incorporation of reliability requirements.

FORM M tI-H-Z REV 11-67

E-22 db Space Division ' North Amencan Rockwell

SP.ACF DIVISION NOR I iI \MFRI(AN ROCK\\ ELI, CORPOR 'TION 12211 I AKEWOOD BOULEVARD * DOWNEY, CALIFORNIA 90241 CODE [DENT. NO. 03953 NUMBER LETE M0823-0007 F ET 1PG I II I I PAGE i

4. QUALITY ASSURANCE PROVISIONS

4.1 Quality Assurance Program. The subcontractor shall have or establish a quality assurance program that conforms to the requirements of Specification STO02GTOo01. Where the subcontractor has an existing program that meets the intent of Specification STOBO2GTOOo1, he may submit this for approval. The subcontractor shall be responsible for all qualification and acceptance tests to demonstrate the ability of the equipment to meet the performance requirements at the prescribed environmental conditions. Every effort shall be made to select equipment which has been previously qualified for a similar application. Where adequate qualification data are not available, the subcontractor shall conduct tests as required to develop and qualify the equipment. Where equipment has been qualified t0 some but not all of the RMU environments, a delta qualification program may be conducted for those particular environments.

4.2 Test Program. The subcontractor shall prepare a test program for the development, qualification and/or acceptance testing of the equipment. The program shall be submitted to NH/SD for review and approval at least 10 days prior to start of any testing.

4.3 Traceability. Traceability requirements for equipment items shall be identified on the SOD and defined in Specification ST0201MA0007.

4.4 Classification of Tests. The tests specified herein shall be classified as follows:

(a) Acceptance tests

(b) Qualification tests 4.5 Definition of Test Classifications.

4.5.1 Acceptance Tests. Prior to delivery and as a condition of acceptance, the subcontractor shall subject each component/subsjstem to the following minimum acceptance tests in the sequence listed unless otherwise specified in the procurement specifica­ tion. Those components/subsystems that are delivered as flight hardware will be subjected to qualification level testing after being installed in the spacecraft, and prior to launch. The equipment so tested shall be capable of meeting the specified performance and life requirements in orbit after being subjected to the actual launch environments. Those components/subsystems that are delivered as flight spares shall be capable of being installed in the spacecraft at any point during its test program, receive the remainder of the qualification tests to which the spacecraft will be sub­ jected prior to launch, and still meet the specified performance and life requirements in orbit.

FORM M 131-H-2 REV 11-67

E-23 9lk Space Division North American Rockwell

S I'A\( DIVISION No Ill AMFI.I\AN ROCEKI I1.I. CORPORAL I1oN 1221 LAIKEWOO BOULEVARO . DOWNEY, CALIFORNIA 90241 CODE IDENT. NO. 03953

NUMBER o823-ooo7 [ REViSION LETTER -

4.5.2 Qualification Tests. The subcontractor shali conduct the qualification tests specified herein on a representative test specimen or shall produce documentation veri­ fying that the tests have been previously performed for another program. Vhere previous­ tests have demonstrated the required performance after being subjected to somej but not all, of the listed environments, a delta ,qualification program shall be submitted to fR/8D for approval. Qualification method for the delta qualification program such as analysis, similarity or test, will be assessed on an individual basis. Qualification tests shall be performed in the sequence listed.

4.6 Test Conditions. Unless otherwise specified for individual tests, all te sts shall be conducted under the following ambient conditions:

.(a) Temperature: 72 t 10F

(b) Relative Humidity: Less than 50 percent

(c) Barometric Pressure: 28 to 32 in Hg correct performance data to 760 nm (29.92 in Hg)

(d) Cleanliness: Less than 1 percent of all particles shall be in excess of 0.5 micron.

I.7 Test Levels. Electrical parameaers of .frequency.voltage-.impedance,.crrent,. pulse timing, and waveform at the equipment Interfaces shall be varied over the fanges specified by the specific SCD.

4.8 Measurement Tolerances- The maximum tolerances allowable in measurement shall be as follows unless otherwise specified in the specific SOD:

(a) Weight: + 0.1% or ' .01 lb, whichever is greater.

(b) Temperature: 1 °F (on controlled temperature sensors and exclusive of instrument accuracy)

(c) Humidity: + 0, - 5% relative humidity

(d) Vibration Amplitude - Sinusoidal: - 10%

Random (Overall EMS Level): + 10% Random (Power Spectral Density): t 3db

(e) Vibration Frequency: - 2% or 1 H. (whichever is greater)

(f) Acceleration: -+04, -5%

(g) Mechanical Shock: Response Spectrum + 5%, - 10%

FORM M 131..H.l REV 1--67

-24 1di Space Division 7'&V NorthAmencan Rockwell

FMC(E DIVISION NOR 111 AMIRIC \N RO( K\ P'1.1.CORPORVIO)N 12214 LAKEWOOD BOULEVARO - DOWNEY, CALIFORNIA 90241 CODE IDENT. NO. 03953 NUMBER MC823-0007 L RPAGE 16

4.9 Test Description and Sequence. The following series of tests will be required for both qualification and acceptance testing unless otherwise specified in the specific SOD. Where test levels and/or test duration differ for acceptance and qualification testing, both requirements are specified.

4.9.1 Examination of Product. The test specimen shall be inspected to verify that the materials, design, construction, dimensions, weight, markings, and workmanship comply with the requirements specified herein and in the specific SOD. 4.9.2 Performance Test. The purpose of this test is to verify performance during and after exposure to the specified environments. The test shall be conducted prior to each of the environmental tests to verify the performance capability of the equipment. The initial test will establish a performance baseline against which subsequent per­ formance tests can be compared. The list of the individual performance tests is specified in the specific SCD.

4.9.3 Leak Detection. Hermetically sealed components and/or pressure containing devices shall be subjected to a leak test as specified in the specific SCD.

4.9.4 Structural Dynamic Tests. The following tests shall be performed to the levels and for the time duration as specified. The test fixture to support the equip­ ment under test should be extremely rigid. The lowest resonant frequency of the test fixture should be well above the major structural resonances of the equipment/test fixture combination. A resonant frequency survey test shall be conducted on the test fixtures prior to the equipment tests.

4.9.4.1 Sinusoidal Vibration. This portion of the test shall be conducted by sweeping the applied vibration once through each frequency range specified in the following table. The level and sweep rate shall be as appropriate for acceptance or qualification test. The test shall be performed once for each of three mutually perpendicular axes.

ACCEPTANCE TEST QUALIFICATION TEST

FREQUENCY LEVEL "g' SWEEP PATE LEVEL "g" SWEEP RATE RANGE (Hz) (o TO PEAK) (OCTAVES/MIN) (o To PEAK) (OCTAVES/Mm)

5-11 0.32 inch 4.o 0.48 inch 2.0 double double amplitude amplitude

11-17 - 2.Og 4.0 - 3.Og 2.0

17-23 4.0g 3.0 + 6.Og 1.5 + + 23-200 - 1.5g 4.0 - 2.3g 2.0

200-2000 ± 3.3 4.0 5.Og 2.0

FORM M t31-H-2 RIEV 1-6

E-2 5 'ft Space Division Nordi American Rockwell

SP.\(E DIVISION NOR]'I \MI:RIC \N ROCKWI:I OROR \VIION 12214 LAKEWOODBOULEVARD DOWNEY,CALIFORNIA 90241 CODE IDENT. NO. 03953 NMBE C823-0007 IRVSNLET PG 17 1

4.9.4.2 Random Vibration. The component/subsystem shall be subjected to a random vibration test in accordance with the following table. The levels and duration shall be as appropriate for acceptance or qualification test. The test shall be performed once for each of the three orthogonal axes:

ACCEPTANCE TEST QUALIFICATION TEST POWER SPECTRAL POWER SPECTRAL DENSITY LEVEL ACCEL. DENSITY LEVEL ACCEL. FREQUENCY 2 RANGE (Hz) (g2 /H,) (g-rms) DURATION (g /H,) (g-rms) DURATION

20-300 0.0013 to 0.02 0.0029 to o1o45 Increasing from 6.1 2 Min. Increasing from 9.1 4 Min. each 20 Hz at a rate each 20 Hz at a rate of 3 db/octave axis of 3 db/octave axis

300-2000 0.02 O.045

4.9.4.3 Acoustic Noise. An acoustic noise test may be substituted for the random vibration test. In this case, the component is not rigidly attached, but is freely suspended in an acoustic test chamber. This test may be desirable, for example, for components having large surface to mass ratios, such as solar arrays. The acoustic levels are to be in accordance with the following table . The test level shallibe as appropriate for acceptance or qualification test. The test-duration shall be one minute for complete exposure.

2 OCTAVE BAND SOUND PRESSURE LEVEL (db, REF. 0.0002 DYIES/Cm CENYER FREQUENCY ACCEPTANCE TEST LEVEL QUALIFICATION TEST LEVEL

15.8 123 127 31.5 123 127 63 125 129 125 130 1311 250 133 137 500 134 138 1000 134 138 2000 131 135 4000 127 131 8000 121 125

OVERALL 140 144

FORM 131-H-2 REV 11-67

E-26 Alk Space Division Vj NorthAmencan Rockwell

SAEDIVISION NOR1 A.MERI\ N ROCKW'EI.I. CORPOR \TION 12214 LAKEWOODBOULEVARD - DOWNEY, CALIFORNIA 90241 CODE IDENT. NO. 03953

1 NUMBER - REVISON LETTER MI8-002307 PAGE 18

4.9.5 Acceleration - Steady State. Components shall be subjected to an accelera­ tion test of one minute duration along each of three orthogonal axes. The maximum steady state acceleration occurs during the firing of the apogee motor. The spacecraft is spin stabilized during this portion of flight. The test level will be established as a resultant of the axial thrust of the apogee motor and the radial acceleration due to spin. The resultant acceleration level shall be computed as follows:

ACCEPTANCE TEST LEVEL QUALIfCATION TEST LEVEL

2 -/(13.97)2 + (0.0000284 rl2) (20.96)2+(0.0000426

r = radius in inches from the spin axis of the spacecraft to the CG

of the component. If r is not specified use 25.5 inches.

N = maximum spin rate expected in RPM (Note: N = 60 RPM)

4.9.6 Thermal Vacuum. The test specimen shall be supported at locations that will minimize the thermal influence of the support during the test. The test specimen shall be placed in the test chamber in such a manner that it is not exposed to any abnormally hot or cold sources. Provisions shall be made for electrical performance tests of the test specimen while in the test chamber as well as for the required instrumentation monitoring. The following criteria shall apply when selecting the test temperatures to be used during the high and low temperature soaks described in subse­ quent paragraphs:

(a) Acceptance Test: Utilize the predicated maximum or minimum temperatures as specified in the specific SCD for operating or non-operating conditions as appropriate. "(b) Qualification Test: Utilize temperature extremes 18 degrees F above the predicted maximmn or 18 degrees F below the predicted minimum temperatures as specified in the specific SCD for operating or nonoperating conditions as appropriate. 4.9.6.1 Corona Test. With the component in operational status for launch, the - chamber shall be evacuated to I x 10 mm Hg. Corona and other high voltage effects shall be monitored throughout the evacuation period.

4.9.6.2 High Temperature Soak. Immediately following the corona check, the component shall be turned off and the chamber shall be evacuated to a pressure - 5 of less than 1 x 10 mm Hg. The temperature of the component shall be increased to the maxi­ mum non-operating test temperature. The component shall be maintained at this temperature for a period of 10 hours after temperature stabilization has been achieved. The component temperature shall then be lowered to the maximum operating test tempera­ ture. The component shall be turned on and the performance monitored for an exposure period of 20 hours after temperature stabilization has been achieved.

FORM M I3-H-Z REV 11-67

E-27 ' Space Division North Ameican Rockwell

SP (;1 DIVIS!ON N(M)II ] .\.II:RI( AN ROCKVrI I. tORPOR X1 I(N 221., LAKEWOODBOULEVARD DOWNEY, CALIFORNIA 90241 NUMBER CODEI, IDENT.,IEiS NO 03953,L TT,I 1 mc823-0007 .1 1AE

4.9.6.3 Low Temperature Soak. At the conclusion of the high temperature exposure the component temperature shall be lowered to the minimum operating test temperature. The component shall remain on and the performance monitored during the temperature transition and for an exposure period of 20 hours after temperature stabilization has been achieved. The component shall then be turned off and the temperature lowered to the minimum non-operating test temperature and maintained at this temperature for an exposure period of 10 hours after temperature stabilization has been achieved. At the completion of the non-operating low temperature exposure, the component tempera­ ture shall be increased to the minimum operating test temperature. The component shall be turned on and the performance monitored for an exposure period of 5 hours after temperature stabilization has been achieved. The chamber pressure shall be maintained at 1 x 10-5 mm Tg or less throughout the temperature excursions.

4.9.7 Electromagnetic Field Measurement. Electromagnetic compatibility and susceptability tests shall be performed as required by the specific SCD for both acceptance testing and qualification testing.

FORM M 131 -H-2 REV 11-67

E-28 Space Division 9D NorthAmencan Rockwell

SPACE DIVISION NORI II AMERIC \N ROCKX ElI. OR ORATION 12214 LAKEWOOD BOULEVARO . DOWNEY, CALIFORNIA 90241 CODE IDENT. NO. 03953 NUMBER 1 REVSIONj PAGE 20

5. PREPARATION FOR DELIVERY

5.1 General Packaging Requirements. The requirements specified herein govern the preparation of items for shipment to NR/SD facilities and test sites. The methods of preservationpackaging, and packing selected shall adequately protect the item against the specified transportation, handling, and storage environments.

5.1.1 Preservation and Packaging. Preservation and packaging shall be in ac­ cordance with the requirements of Level A of Standard MIL-STD-794.

5.1.2 Packing. Packing shall be in accordance with Level B of Standard MIL­ STD-794.

5.2 Packaging Design Criteria. The requirements specified herein are design requirements only and do not require demonstration.

5.2.1 Size and Weight. The gross weight and the cubic displacement of the package shall be held to a minimum consistent with the requirements of this specifica­ tion.

5.2.2 Drop Height Criteria. 5he package shall be capable of withstanding the conditions described in Paragraph 3..2.1.1-a(1) of this document.

5.3 Marking for Shipment. The container shall be labeled and marked in accor­ dance with Standard MIL-STD-129, and all applicable freight tariffs. It shall include as a minimum. -R/Space Division Control Number

Item Name

Manufacturer Type or Part Number

Quantity in Package- ot Number (if Applicable)

Serial Number (Only if Serialized for Control)

ManufacturelA

Purchase Order Number

Date of Packaging

FORM M 131-H -2 REV 11-67

E-29 Space Division North American Rockwell

SPICE DIVISION NORTH \MERIC\N ROCKIElI. :ORPOR\TION 12214 LAKEWOOD BOULEVARD * DOWNEY.CALIFORNIA 90241 CODE IDENT. NO. 03953 ' ° jNU L MC823-0007 E, SIO" LETTER P 21

6. NoTES

6.i Definitions. 6.1.1 N/SD., NR/SD is identified as,the Space Division o.North American,. Rockwell Corporation, Downey, Californ.ia&.

6.1.2 Subcontractor.- The term'subcontractor shall e the manuficturer' to whom the Purchase Order/Request for Quotation is addressed.

FORM M 131-H-2 REV 11-67

E-3 0 NASA FORMAL REPORT

FFNo 665 Auq 65