Testing the Man-Rated Launch Vehicle

The Space Congress® Proceedings 1966 (3rd) The Challenge of Space

Mar 7th, 8:00 AM

Francis X. Carey Martin Company

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Scholarly Commons Citation Carey, Francis X., "Testing the Man-Rated Launch Vehicle" (1966). The Space Congress® Proceedings. 2. https://commons.erau.edu/space-congress-proceedings/proceedings-1966-3rd/session-6/2

This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. TESTING THE MAN-RATED LAUNCH VEHICLE

Francis X. Carey

Martin Company Cocoa Beach, Florida

Although manned space flight is still in From this mathematical relationship, it its infancy, testing of launch vehicles can be seen that achieving a satisfactory has progressed to a high degree of so level of probability of crew survival phistication. required that careful consideration be given to such launch vehicle items as: As of December 1965, the Martin-built Gemini launch vehicle has launched 1. Component and/or system, redundancy seven Gemini spacecraft successfully which can improve the reliability out of seven attempts. This remarkable of the launch vehicle record was made possible by two facts: 2. Analysis of launch vehicle failure 1. The basic reliability of the modes followed by design of a re hardware liable malfunction detection system or MDS 2. The test program. 3. Functional utilization of the crew This paper briefly describes the Gemini as part of the malfunction detection launch vehicle noting the major differ system ences between it and the Titan II and discusses the test program. It is not 4. Emphasis in launch vehicle checkout proposed that this is the only method on minimizing the possibility of of testing a man-rated launch vehicle; launching a bad vehicle however, it is a successful one. 5. Test, countdown, and launch proced The Gemini launch vehicle is a basic ures that will lead to maximum Air Force Titan II which has been modi probability of launching a good fied in certain areas to achieve man- vehicle rating (see Figure 1). 6. System simplification where possible Launch Vehicle Man-Rating to achieve reliability.

Man-rating is the philosophy and plan Man-rating a Titan II is a many-sided for marshaling the disciplines necess process conceived by Martin and the Air ary to achieve a satisfactory proba Force to improve the reliability of the bility of mission success and crew basic vehicle by modifying existing sys survival. This probability may be tems, by using redundant components, by expressed in terms of reliability of adding special systems for crew safety the malfunction detection and escape purposes, by special handling of critical systems. Mathematically, these terms components, by meticulous selection of are linked in the equation: qualified people, and by developing pro cedures in the entire design-production- = RLV *MDS manufacturing-test-launch cycle that es tablishes as a goal flawless performance where: from the launch vehicle.

PCS = probability of crew survival Titan II Modifications for Gemini

RT XT = reliability of the launch The high reliability of the ultimate man- vehicle rated Gemini-Titan launch vehicle was the made early in the = reliability of the malfunction result of two decisions engineers-- detection and escape systems. program by Martin and Air Force

332 first, to make as few changes as possible with one important exception: ASIS was in the basic Titan II and secondly, to completely automatic (after a serious make every necessary change count toward malfunction was discovered, ASIS auto overall reliability. Each of the modifi matically initiated mission abort), cations added to reliability and safety while MDS only provides vehicle condi and contributed to man-rating. tion information to the astronauts. In other words, MDS signals to the astronaut The major modifications, accepted only the proper or improper functioning of the after careful review by a top level Air launch vehicle and allows him to make the Force-Martin engineering board, were: decision as to whether to continue the mission in the event of a malfunction or 1. The addition of a malfunction de when to abort the mission if this should tection system designed to sense become necessary. Only in the event of problems in any of the vital an engine hardover condition does the MDS booster systems and transmit this automatically act and in this case, only information to the astronaut crew to switch over to redundant hydraulic, flight control, and electrical systems, 2. A redundant flight control system not to abort the mission. As a result, which could take over should the the MDS is a relatively simple system primary system fail in flight with a high degree of reliability. It takes advantage of one of the lessons of 3. Redundancy in the electrical sys Project Mercury - that man can function tem with necessary changes to in space as a working pilot, not just as provide power for such added a passenger in his spacecraft. The MDS, launch vehicle equipment as the which is completely redundant, monitors MDS during flight: propellant tank pressures, staging of the launch vehicle, thrust 4. Substitution of radio guidance chamber pressures, electrical system volt for inertial guidance used in the age, and turning rates which would indi Titan II ICBM version to provide cate a need for action if the launch a weight reduction and also to vehicle structural limits are approached. provide a more responsive system during critical orbital injection Hydraulic System

5. Redundancy in the hydraulic sys The launch vehicle's hydraulic systems tems where desirable for pilot control the position of the Stage I and sa±£ry, sucn as hydraulic actua Stage II rocket engine thrust chambers tors for engine gimbaling in response to electrical signals from the flight control system. Changing the 6. Instrumentation to provide addi thrust direction in this manner, the tional data during pre-flight launch vehicle is steered along the de checkout and flight not considered sired line-of-flight by making corrections necessary in the ICBM version. in pitch or yaw axes. The thrust chambers are gimbal-mounted to allow two degrees of Malfunction Detection System freedom. Stage I roll control is provided also by varying direction of the two The malfunction detection system (MDS) thrust chambers. A separate roll control (Figure 2) is perhaps the most signifi nozzle with one degree of freedom is pro cant modification made to Titan II to vided in Stage II which has one thrust prepare it for the manned Gemini mis chamber. sion. The MDS monitors operation of vital launch vehicle subsystems and The Gemini-Titan launch vehicle has three signals the spacecraft crew if a mal separate hydraulic systems - the Stage I function takes place. The Gemini-Titan primary system, the Stage I secondary sys MDS is comparable in function to the tem, and the Stage II system. A schematic launch vehicle ASIS (Abort Sensing and of the redundant Stage I hydraulic system Implementation System) of Project Mercury is shown in Figure 3.

333 Stage I Primary, Basic components chamber for pitch and yaw control, and of the Stage I primary hydraulic system one on :he off-center turbine exhaust are an engine-driven pump, an accumula nozzle for roll control. tor/reservoir, four servo actuators, and an electric motor-driven pump. An axial piston, pressure-compensated, variable volume engine-driven pump The Gemini-Titan flight control system pressurizes the system at 3000 psi constitutes a basic change from the during Stage I flight. The pump is Titan ICBM system. Specifically, the driven by the accessory drive pad on system is designed to withstand any the turbodrive assembly of the Stage I single malfunction (failure) and com engine. The servo actuators control plete the launch phase successfully. the movement of the Stage I thrust The design achieves this increased re chambers. The actuators are built as liability by complete flight control tandem units consisting of two complete system redundancy (Figure 4) . electro-hydraulic servo system sections. One servo section of each actuator is A complete secondary or redundant flight connected to the primary hydraulic sys control system is provided in the Gemini- tem; the other servo actuator section Titan launch vehicle to protect the in each actuator comprising a portion astronaut crew from this one event. of a secondary or redundant hydraulic Switchover from the primary to second system* The actuator sections are in ary system in the event of an engine dependent but interconnected by a hardover condition is automatic and special "switchover"valve that permits requires only 15 milliseconds. The only one of the servo sections to be switchover is final and no provision operable at any given time. is made for switching back during Stage I flight. After staging, the pilot may Stage I Secondary. The secondary switch back to the primary mode. or redundant Stage I hydraulic system Switchover to the secondary system is identical to the primary system. It also can be effected by three other operates in line with the secondary or methods: redundant flight control system. The same electric motor-driven pump is used 1. Astronaut command for both secondary and primary systems during checkout through use of a system 2. Vehicle overrate detected by the test selector valve. The secondary sys MDS rate sensors tem is pressurized throughout flight, although not in use unless switchover 3. Loss of Stage I primary hydraulic occurs. The actuator switchover valves system pressure. are designed to sense primary system pressure and initiate switchover to the The primary flight control system con secondary hydraulic system in case of sists of the following: any failure within the primary system that would result in a loss or degrada 1. Three Axis Reference System (TARS). tion of pressure below a predetermined Flight-proven and adapted from the value. Titan I ICBM program, the TARS is located in the equipment bay between Stage II. The Stage II hydraulic the tanks of Stage II. It provides system is not redundant since studies information concerning angular dis of potential malfunctions did not indi placement along the three perpen cate a necessity for redundancy such as dicular axes of roll, pitch, and is the case in the event of an engine yaw by use of gyroscopes. Included hardover condition at Max Q in Stage I. in the TARS unit is a programmer The components of the Stage II system which, as a function of time during are basically the same as for the Stage Stage I flight, changes the angular I systems except that three actuators reference of the launch vehicle in are used - two on the single thrust the pitch and roll axes, thereby

334 initiating changes in the direction ponents of the launch vehicle's of flight along these two axes angular rate during Stage I flight. according to a preplanned flight The rate gyro output signals are trajectory. In other words, it supplied to the autopilot. serves a guidance function in Stage I flight signaling commands to such 4. Autopilot. The autopilot is located other units as hydraulics by way of between the Stage II tanks and is the adapter package and autopilot. modified from the Titan ICBM auto The roll programmer can be continu pilot only where necessary to meet ally updated during the prelaunch specific Gemini mission requirements. countdown to set a new azimuth, Included in the autopilot are: three such as might be necessary in a axes rate gyros for Stage II flight; rendezvous mission to achieve the an 800 cycle static inverter provid correct orbital plane. The pitch ing magnetic amplifier and rate gyro program is not changed during power supply for both Stage I and countdown since it is fixed ahead Stage II rate gyros; and circuitry of time for each individual necessary to accept signals from mission* During Stage II flight, Stage I and Stage II rate gyros and the TARS accepts steering signals to amplify, distribute, and condition or angular reference changes in attitude reference signals from the pitch and yaw from the radio guid adapter package to the hydraulic ance system and executes these servo actuators. commands by again signaling other components of the flight control 5. Mod III G Radio Guidance. The radio system. The TARS also actuates guidance system, discussed in detail auxiliary switches in accordance in another section, provides pitch with a preset time program to pro and yaw guidance signals to the TARS vide such functions as arming of through the adapter package during the sensors which signal stage Stage II flight. separation at the proper point in flight. 6. The secondary or redundant flight control system consists of the Adapter Package. The adapter pack following components: (1) Stage I age, also located between the Stage rate gyro: a duplicate of the pri II tanks, conditions the attitude mary system unit (2) Autopilot: outputs of the TARS for input in primary system duplicate. the autopilot. At the proper time as programmed in the TARS, the 7. Spacecraft Inertial Guidance System. adapter allows the pitch and yaw The spacecraft IGS provides those steering signals from the radio attitude stabilization signals to guidance system to be applied to the secondary system autopilot which the TARS guidance amplifiers. The would have been provided during Stage adapter package also houses the I flight in the primary system by the switchover relays (redundant) which TARS and during Stage II flight by effect the primary-secondary flight the radio guidance system. The iner- control system shift. During check tial guidance system is pre-programmed out of the booster, performance of to perform this function in the event all fifteen flight control system of switchover to the secondary system. gyros are monitored through sig nals amplified by the adapter pack Gemini Test Program age. The test program on the Gemini launch vehi Stage I Rate Gyro. Located in the cle is a repetitive series of detailed, inter-stage section, the Stage I quantitative tests starting at the Vertical rate gyro is a Titan II flight con Test Fixture in the Martin Company's manu trol unit. It contains three gyros facturing plant at Baltimore, Maryland and to measure pitch, roll, and yaw com culminating in the launch from Complex 19

335 the responsible design engineers at Cape Kennedy, Florida. viewed by for specification compliance and entered (see Figure 6). In developing this program, great care on the trend charts taken to have the tests and proced was subsystems have been ures at both facilities as nearly iden When the individual the 235 airborne telem tical as possible. The aerospace ground verified, each of is calibrated - usually equipment, with the exception of the etry measurements - and data reduction propellant loading system, at both fa four to six points cilities are identical. curves are drawn. the subsystem test se The data gathered in the Vertical Test The next step in and analyzed by the quence is the telemetry ambient test. Fixture is plotted exer to verify initial per Here, the individual subsystems are system designer sequence Each time a test is repeated, cised in a tightly-controlled formance. system is on compared with the original. while the airborne telemetry the data is this data data accompanies the and radiating and recordings of All this test reduced and entered to Martin's Canaveral Division are made. The data is vehicle Data Book for future at Cape Kennedy where the process is in the launch vehicle as a data reduction tool. continued. reference and use first major system test Factory Test and Vehicle Acceptance Systems Tesj:. The which the launch vehicle undergoes is the Test. This sequence of testing and acceptance Acceptance Mode Verification The to the Combined Sys the Vertical Test Fixture is shown test is quite similar at and the combined system Figure 5. A brief description of tem Acceptance Test in in that a short minus these activities follows. tests at Cape Kennedy time countdown is conducted, followed by a and Alignment. simulated flight. This data is again re Launch Vehicle Erection and used as a The vehicle is erected and checked for viewed by the design groups launch vehicle. vertical alignment structural twist and baseline for the levelness of the gyro mounting pads. Combined System Acceptance Test. The Com Test specifies in Post-Erection Inspection. Approximately bined System Acceptance for a complete detail the steps necessary to bring the one week is allotted equip of the vehicle launch vehicle and aerospace ground physical inspection condition to cabling, and black box ment from a static power-off structure, In addition, installations prior to initiating the a countdown configuration. for performing an subsystem test phase. it gives direction abbreviated count and two simulated flights section on Cape Kennedy Subsystem Functional Verification. (for details, see Verification of the launch vehicle sub Tests). with voltage standing systems commences data from and attenuation tests of the Launch Vehicle Acceptance. The wave ratio Test and and radio frequency cabling the Combined System Acceptance waveguide test data are pre resistance checks of the auxiliary all manufacturing and and to the Vehicle Accept and instrumentation busses. Power is pared and presented management and engin then applied to the individual busses ance Team by Martin for their review. The and systems sequentially. eering personnel Vehicle Acceptance Team is chaired by a Space vehicle subsystems, i. e., senior officer of the Air Force The launch Office flight controls, tracking Systems Division Gemini Program electrical, of manage frequency, propulsion, mal in Los Angeles and is composed and radio from the function detection system, and instru ment and technical specialists are now subjected to the Air Force, NASA, and Aerospace Corporation. mentation Team determines, first of the series of meticulous tests When the Vehicle Acceptance of the launch vehi which finally culminate in the launch through its own analysis that the launch from Cape Kennedy. All data is re cle history and test data,

336 vehicle meets the stringent requirements detection system; and a simulated flight of the Gemini program, it is accepted by terminating at sustainer engine cutoff. the Air Force and prepared for shipment An electronic spacecraft simulator is to Cape Kennedy. used to provide proper electrical loads and simulated guidance signals to the Cape Kennedy Tests launch venxci.e. All systems are then recycled to T-3 minutes and another The pre-launch testing at Cape Kennedy countdown ensues. During the plus time and the Air Force Eastern Test Range can run of this second simulated flight, the be divided into two phases: launch vehicle remains in the primary guidance and flight control mode. Veri 1. Launch vehicle fication of correct response to the pitch and roll program from the Three Axis 2. Integrated launch vehicle and Reference System as well as steering spacecraft. commands from the radio guidance system and system response to discrete commands, Launch Vehicle i. e., gain change, staging arm, booster engine cutoff, and sustainer engine cut Subsystem, The first portion of off is made. The launch vehicle systems the testing at the Cape is quite simi are again recycled to a T-3 minute pre- lar to the testing at the Vertical Test launch configuration for a third and Fixture in Baltimore (see Figure 7), final simulated flight. During the plus However, in place of the subsystem time run for this test, all launch functionals, vehicle an abbreviated subsystem umbilicals are removed in their normal retest has been substituted. This sequence. At the normal staging time, abridged testing in no way detracts booster engine cutoff, the electrical from our confidence in launch vehicle connectors at the staging interface performance. Rather, it is made possi plane are disconnected. The launch vehi ble by our test philosophy which dic cle is evaluated not only for response tated that all test data remained with to proper commands but also to ensure the vehicle until launch. It is there that there is no improper subsystem fore possible for the data recorded at interaction. the Cape to be compared with the acceptance data and analyzed for trends Combined Launch Vehicle and Spacecraft indicative of an incipient failure. This analysis is performed concurrently Most of the combined launch vehicle and by the Martin Company Project Engineer spacecraft testing is done on a system ing Section at the Canaveral Division basis. However, since the inertial guid and by the various design groups at the ance system is used as the secondary Baltimore Division. It should be noted launch vehicle guidance system, subsystem that this data comparison and trend testing, i. e., detailed quantitative analysis do not cease when the individ verification of system gain and response ual subsystems have been reverified. is performed. Likewise, since the primary The subsystem retest data is added to purpose of the malfunction detection system the vehicle history and compared with is to inform the pilots of vehicle per the data gathered in all the combined formance, this subsystem also undergoes tests which follow (Figure 8). more detail verification of the interface.

System. The launch vehicle is Electrical Interface Integrated Vali declared ready to accept the spacecraft dation. All interface wiring across the upon the successful completion of the launch vehicle/spacecraft interface is Pre-mate Combined System Test. The Pre- redundant. To ensure that this is true, mate Combined System Test consists of a each circuit is subjected to a power-off countdown from T-45 minutes through T~0; resistance test with first one and then a simulated liftoff; switchover to the the other electrical interface connector secondary guidance and flight control connected. When verification of the proper system initiated by the malfunction resistance on all circuits is complete.

337 both connectors are mated and the com system performs properly by making de plete launch vehicle/spacecraft is ready tailed, quantitative measurements of the for functional test. Some of the func attitude and rate gains. Proper system tional tests which the space system is phasing is determined by placing the subjected to and a brief description of inertial platform in the inertial mode each follows. and utilizing the earth's rotation to generate an output to the flight control Liftoff and Test Conductor Abort. system. Verifies that when the launch vehicle lifts off the spacecraft receives the Joint Combined Systems Test proper indication. Verifies that the test conductor abort command is re The Joint Combined Systems Test is the ceived by the spacecraft by both radio first major exercise of the complete frequency and landline methods. space system and is designed to verify compatibility of the launch vehicle and Ascent Simulation. The Ascent Sim spacecraft in a simulated flight config ulation is quite similar to the first uration. This test consists of: run of the Pre-mate Combined System Test. However, the spacecraft inertial 1. An abbreviated countdown and liftoff guidance system now furnishes the pitch followed by a simulated abort initia and roll programs, certain steering ted by the pi lot commands, and discrete commands to the launch vehicle. Switchover to the 2. The launch vehicle and spacecraft secondary guidance and control system are then recycled to T-45 minutes is initiated by the pilot. for a countdown and simulated flight utilizing the primary guidance and Fade-in Demonstration. Verifies flight control system. that when the space system switches ftom primary to secondary guidance and Simulated Flight Test control the inertial guidance system computer recognizes the amount of atti The Simulated Flight Test is the final tude error and corrects the flight path exercise of the complete launch vehicle/ over a period of time rather than apply spacecraft system prior to initiating the ing a step function. launch countdown* This test consists of three simulated countdowns and flights Fuel and Oxidizer Tank Pressure which are identical to tests previously Meter Calibration, A three-point cali described, i. e,, abort, secondary guid bration and comparison of the malfunc ance and control, and primary guidance tion detection system propellant tank and control. Immediately following the pressure transducers and spacecraft Simulated Flight Test, the launch vehicle analog meters. is declared ready for flight and special quality and security procedures are put Primary to Secondary Switchover in effect to ensure system integrity is Test, Verifies that each of the hy maintained. draulic inputs to the malfunction detection system will initiate switch Launch Countdown over independently. Since the launch vehicle and spacecraft Staging Interface, Verifies the are both relatively complex systems and spacecraft receives proper indication independent during normal flight, a split of launch vehicle staging. countdown was developed for Gemini (see Figure 9). The split count allows for Joint Guidance and Controls Test maximum flexibility and minimum inter dependence between the two systems during The Joint Guidance and Controls Test the countdown. assures that the launch vehicle/space craft secondary guidance and control The launch count commences on F-l day

338 approximately nineteen hours prior to engineering changes, production, launch. In inspec a four-hour period, all of tion, testing, handling, acceptance the launch and vehicle/spacecraft interface launch; emphasis on configuration is reverified. docu The count is then held mentation and verification control; until T-360 minutes and on launch day. It extensive data and procedural reviews. is during this hold that the launch vehicle is loaded with propellants and As part of the pilot safety program, final topping of some the of the spacecraft Air Force imposes stringent requirements cryogenics is accomplished. during the acceptance phase. Hardware is not accepted until The Range count the Air Force is con is initiated at T-240 vinced that the hardware minutes with and documenta the spacecraft and launch tion comply with appropriate vehicle counting independently. specifica It tions and other contractual requirements should be noted that one of the major and meet the requirements for the Gemini precepts of the Gemini test philosophy mission. Acceptance is characterized by dictated that the launch vehicle be a methodical completely approach and an uncompro verified as ready for flight mising attitude. prior to crew ingress (T-95 minutes). Therefore, there are no tests performed In discussing the pilot safety program, after ingress that have not been per this paper is confined to its influence formed earlier in the count. and impact on the pre-launch test program. After crew ingress, the countdown - Test Procedure Control with the exception of the mechanical functions necessary to lower the erec The Gemini test procedures used tor and prepare at Cape the stand for launch - Kennedy are written by the Canaveral is an exact replica of the count per Division of the Martin Company. These formed for the Pre-mate Combined Systems basic procedures are prepared by the Test, Joint Combined Systems Test, and engineers assigned to the Launch Opera Simulated Flight Test, tions Section and are governed by con tractual and engineering specifications. For a closer insight into the problems The Gemini test procedures associated are written with planning and conducting in considerable detail, a countdown, leaving nothing see "Rendezvous Launch Op to the discretion of the operator. erations Planning."^ The specification revision designation is noted on the draft.

The procedure is then forwarded to A launch vehicle pilot safety program Configuration Management where has been established it is by the Air Force verified that the 'written Space to 1 level of Systems Division to ensure that the referenced specifications a concern is the for safety is manifested in latest released engineering. This plans, reflected is in appropriate activity, performed using a computer adequately documented, tab run and thoroughly which is updated daily. assessed prior to launch. The program is implemented in two ways. First, the The procedure next is program routed to the ensures a continuous monitoring Gemini Project Engineering effort commencing with and Quality the preliminary Engineering sections for reverification design and continuing through launch. of technical adequacy, specification Second, the program concentrates con compliance, and adherence to Martin siderable effort at key focal points quality standards. when major problems arise. Assurance that nothing has been neglected is pro The Pilot Safety engineer reviews the vided by following a pattern of rigorous procedure for compliance with the total technical monitoring of associate con pilot safety program. tractors' activity; rigid control of all phases of design, development, The procedure is then forwarded to the

339 customer, i. e., the 6555th Aerospace flight development effort. Test Wing, for review and coordination.

After all approvals are secured and all agencies satisfied, the procedure is Another aspect of the pilot safety program - published and released for use (Figure one which cannot be over-emphasized - is 10). the motivation and skill of the launch crew. Each and every technician and engin Now, the typical procedure .starts through eer working on the program must be aware of the test cycle. The Official Test Copy his responsibility for pilot safety. For (OTC) is issued to the responsible sys in the final analysis, all of the manage tem engineer in accordance with the test ment effort, data reviews, trend analysis, schedule. Attached to the Official Test and basic vehicle reliability can be ne Copy are several Procedure History Sheets gated by an undetected human error „ for documenting all pertinent facts, deviations, and anomalies concerning the The average man assigned to the Gemini. test. The complete package is bound and launch team has over seven years of sealed by the Quality Section to assure Martin launch experience. Most of this that all pages remain in the Official experience was on the earlier Titan I Test Copy. and Titan II vehicles. Nevertheless, upon transfer to the Gemini program, After the completion of the procedure, they were again subjected to rigorous a post-test critique meeting is held classroom training, written and oral during which all recorded data and Pro exams, and performance demonstrations cedure History Sheet items are reviewed before being certified as team members. and accepted or rejected. Present at This training has continued throughout this meeting and responsible for verify the program, ing that the data is satisfactory and the system is performing properly, are Methods of motivating people to assure the Operations engineer and the appro maximum quality vary from the classical priate system engineers from the Project poster "Uncle Sam Wants You" to pep talks Engineering and Quality Engineering and lectures. The ones which seem to sections. have worked on Gemini are the sense of identification with the program and the The test procedure and all associated pride in a job well done. These have data, i. e., strip charts or other re been greatly enhanced by the almost daily cordings, are then forwarded to the contact with the flight crews, the estab Quality data storage area for review lishment of many space records for the by the Pilot Safety Working Team. United States, and the many words and letters of commendation from the Air The Pilot Safety Working Team consists Force, NASA, and the 'guys who drive of representatives of Martin Company them 1 . Engineering Section; Customer Quality7 Aerospace Corporation; and the 6555th versus Program Needs Aerospace Test Wing. They conduct a separate review to assure completeness The confidence level necessary to commit of the documentation and procedure. to a manned launch is achieved increment- ally during the test program for each It is readily recognizable that these launch. Affecting this 'confidence' are additional reviews and reverification the results of the previous f lights f the of the documentation add considerable Vertical Test Fixture tests, and Eastern cost to a program over and above what Test Range tests, the data and trend analy would normally be expected in a weapon sis, and the performance of the launch system development program. From a crew. launch operations standpoint, it can be estimated this cost as approxi Assuming that this is true, the question mately thirty per cent over a normal naturally arises as to how the Gemini 7

340 and 6 mission was possible. The answer is simple. The Gemini 7 and 6 mission was possible because of the Gemini test philosophy.

GT-6 was completely checked out accord ing to the normal Gemini schedule, in cluding a launch countdown which termi nated prematurely due to a target vehicle problem, GT-6 had performed perfectly during all phases of the ground test and analysis of the test data showed no trends indicative of a potential problem. Extreme care was exercised in removing it from the launch complex. Only those electrical connectors at the staging and spacecraft interface planes were discon nected, making the retest minimal.

After re-erection, the individual sub systems were reverified, using somewhat abbreviated procedures prior to the Simulated Flight Test. From the Simu lated Flight Test through to launch, normal procedures were followed. The data from the Simulated Flight Test was compared by use of a computer with the Simulated Flight Test performed prior to the first launch attempt and found to be an overlay. The launch vehicle was again performing perfectly according to the engineering criteria and the emotional criteria of confidence was satisfied.

Although the Gemini test program is con trolled by very rigid application of the engineering and quality disciplines as outlined in the procedural control, data and trend analysis, and pilot safety program, it has maintained the flexibility to respond to changing pro gram needs.

References:

1. Rendezvous Launch Operations Planning by Mark Goodkind, Proceedings of Third Space Congress

341 Rendezvous guidance & recovery system

Re-entry capsule Spacecraft Adapter section

Separation point Oxidizer tank

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'Equipment bay contains: Oxidizer Batteries tank Malfunction detection system IMDS) units Range safety command control system Programmer Three-axis reference system (TARS) Radio guidance system (RGS) Autopilot Instrumentation and telemetry system

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-0.011

-089

+0.025

+0.014

(ETR

DATA

5750

Note

effective

9-20-65

(Disconti

25.67

109.65

EIIV

No.

144.05

162.25

N/A N/A

N/A

-089

N/A

TREND

(See

Date Test

Tests

VTF

tolerance

011/01

6-25-65

and

number

GLV-_1_

CSAT

No.

25.66

104.88

317.03

-0.060

+0.117

144.53

162.33

-0.062 -0.116

+0.088

-0.085

+0.025

+0.015

-0.025

-0,010

-089

was

Date

Test

time

New.

Ref.

L0±

seconds,

T±.

Time

relay

configuration

1,22

»/B

"/s

»/E

delay

Value

sec.

.133

sec.

sec*

-.086

;

-.107

-.047

-.020

27.17

time

10/25

110,00

Tolerance

Spec

104.96

1.47

1.07 144.64

317.44 3.19

162.56

1.66

and

«•

Nominal

24.83to

Volts

-.042

-.059

-.095

-.007

added

-.004

tests

Attempt

517R2

ETR

Shutdown

Ecp

BGC

all

Launch

AGC

Initiate

Inverter

Step

Program Step

Step

Parameter

Program

Discreet

Change

Discreet

Program

CPS

Stage

Stagell

80O

Pitch

Voltage

TARS Gain

TAPS

Arm

TARS

Roll

First

Second

Configuration

Arm

TARS Guidance

Pitch

Third

Pitch

Note

No.

Meaa

0727

0728

0735

0739

0150 0150

0740

0150

0153

0151

0153

0152

NOTES:

No.

Line

10.

EVENTS EVENTS

2233

EVENTS

VEHICLE VEHICLE

22222233 22222233

LEGEND

122222222

LAUNCH LAUNCH

SPACECRAFT SPACECRAFT

23456789012345678901

EVENTS

DECEMBER

11111111112222 11111111112222

01 01

.0.1.2.3.4.5.6.7.8.9.0.1.2.3.4.5.6.7.8.9.0.1

H-5

OF OF

iiQA*ATQo

TP787a.H-452 TP787a.H-452

SERVICING SERVICING

SEQUENCE SEQUENCE

1122222222223

!!!

P778 P778

NOVEMBER

.0.1.2.3.4.5,6,7.8,9,0.1,2,3,4,5,6^7,8^0

CATAPULTS

a a

INTERCOM INTERCOM

& &

INTEGRATED INTEGRATED

PRE-POWER PRE-POWER

1 1

TOlACXTOeJ

ACT.,SEATS ACT.,SEATS

GT-7 GT-7

INSTALLPYRO INSTALLPYRO

& &

APPROVAL

rrri

MINI MINI

MTWT MTWT

rrr rrr

SEATS SEATS

01 01

DIVISION DIVISION

OCT

-10-61

222222233 222222233

IY IY

7-J4-4S 7-J4-4S

rri rri

INSTALL INSTALL HATCH

ii ii

IIVISIDMl IIVISIDMl

ISSUID: ISSUID:

PRIPARID PRIPARID

CANAVIRAL CANAVIRAL

i i

REMOVE REMOVE

1122222222223

89012345678901

MTWT MTWT

B.9.0.1.2.3.4.5.6.7.8.9.0.1

11222 11222

CO CD a s * ^

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I P% I C! AA?- te% ^' -t_ •3'i v i \ *t c^ is •n . o ^ * i i c I >3 ^ : i^ $.. i & n 1 §^ -1 I "IvC • OE *» c i jC ia * * ?.R i:U| 0 r^ 5 J 1' r1>^. ^ 'H>j • r n o' A 1 -iA^ jj , < c A> 5! O^" i 1 ^r'r ^ L 1 U ^ v1 ft ^ I ^ !• ;) : v, § 1 39 J ° -/ ^ $- \ J ssf TA. (J) 3 -1 ^ 1 \

| 1

| | i i t \ |

CANAVERAL DATA ENGINEERING

MEASUREMENT N0._2i22___ RANGE + J-- 2 ? MCHES

DESCRIPTION TRAVEL ACT. NO. 1 PITCH

•

L L

~\ ~\

-J

P P

./irj

PST PST AiCO AiCO

R/G R/G

OPEN

ASC-j

WERED^ WERED^

-H -H

| |

in in

HPST

G G

———— ————

CALS CALS

' '

— —

ID ID

3CT 3CT

P/V P/V Fj Fj

/ /

' '

Yn Yn

\ \

n n

SENSOR SENSOR

PRESS. PRESS.

2 2

/CM /CM

Q[J~\

BANGE BANGE

READOUT-,

^ ^

AUTO AUTO

i i

TEST-P TEST-P

ALL ALL

\ \

Hjrt Hjrt

arc arc

JJMBLENT JJMBLENT

TRANSIT TRANSIT

TANK TANK

P-C-^"^ P-C-^"^

*L *L

L! L!

SSSSOR SSSSOR

POWER— POWER—

LOOP LOOP

FLIGHT FLIGHT

PT PT

n n

,, ,,

COUNTDOWN

OX OX

RESTRICTED RESTRICTED

PIR PIR

PTHAL PTHAL

READOUT READOUT

' '

DESTRUCT DESTRUCT

iLirr-oPF iLirr-oPF

l" l"

| |

A/B A/B

TANK TANK

/ /

——— ———

OPES OPES

' '

OPEN OPEN

RANGE RANGE

i i

S/D, S/D,

GLV GLV

1. 1.

875/ETR 875/ETR

J J

Nov. Nov.

RANGE RANGE

¥// ¥//

y y

^ ^

|— |—

| |

^~1

CONN CONN

1,1 1,1

' '

SILENCE SILENCE

x x

DESTRUCT DESTRUCT

CREW CREW

QQ QQ

—— ——

/ /

RF RF

PCN PCN

22 22

t-ASCO, t-ASCO,

y-^^sr y-^^sr

3 3 zi'y

( (

TT TT

| |

DTTERFACE DTTERFACE

H H

L L

ORD ORD

rf°""\ rf°""\

ihn ihn

- -

* *

. .

__^ __^

' '

S/D

| |

i i

3EACC

CESXX.

ASCO ASCO

\ \

, ,

RAMGE RAMGE

RESTRICTED RESTRICTED

3Arr 3Arr

DESTRUCT^fl

—— ——

START

______

PROCEDURE PROCEDURE

PAGE PAGE

MODEL MODEL

\ \

_ _

.1 \ \

'CALIBRATIONS 'CALIBRATIONS

——— ———

| |

. .

1 1

^^** ^^**

gjjjj

-p. -p.

.VND .VND

^p—— ^p——

TEST—— TEST——

READOUT- READOUT-

iir-iL.^: iir-iL.^:

IOOP IOOP

ONLY ONLY

"- "-

^JBIESTS ^JBIESTS

OPEN OPEN

STRANGE STRANGE

GREW GREW

POWER

lOADIHG lOADIHG

PRESSURE PRESSURE

GROUND GROUND

PPP^nJKT PPP^nJKT

BLANKET BLANKET

READOUT READOUT

PBOPELLANT PBOPELLANT

C C

COUNTDOWN

RANG RANG

| |

]

• •

^^ ^^

- -

PST

i i

MASTER MASTER

nJ nJ

n n

>-S0ITCBDVm: >-S0ITCBDVm:

u u

RISrmiCTBtt RISrmiCTBtt

O-SWTTCHDVER

i i

v v

GCT GCT

READOUT READOUT

nn nn

1 1

Ufc|

SENSOR SENSOR

i i

SHOTDOWM SHOTDOWM

READINESS; READINESS;

— —

GDT GDT

LtfT-OW LtfT-OW

1 1

mmm mmm

RANGE RANGE

^TANK ^TANK

S/C S/C

OVERRATE OVERRATE

n n

^* ^*

J^^cf

^.BEACON ^.BEACON

^^^^**

FAD FAD

CTXA*

SA> SA>

I I

nrTERFACEHJ nrTERFACEHJ

\ \

S/D S/D

ABd^^P ABd^^P

^P-^- ^P-^-

OVlMATBy OVlMATBy

fc fc

———— ————

v*»inrrs v*»inrrs

ON ON

TEST TEST

TEST

TEST TEST

awms

TEST

OH OH

Interface).

TUT

OH OH

PREW, PREW,

STATION STATION

OBBCK

OBSTRUCT OBSTRUCT

(Non- (Non-

ASCENT ASCENT

TEST TEST

TRANSMITTERS TRANSMITTERS

LQADlJfc

TAME TAME

TEST TEST

GUIDANCE GUIDANCE

ON ON

OH OH

and and

RECORDERS

OPERATIONS

ntTElFACE ntTElFACE

FM FM

RECEIVERS RECEIVERS

CARRIER CARRIER

OH OH

TEST TEST

SHUTDOV* SHUTDOV*

CBECK

RF RF

TEST

CUURMNJHJ

GOIHtOLS

1 1

III III

and and

ACCESS

POWER POWER

INTERFACE INTERFACE

HOC HOC

flflClLAHgK, flflClLAHgK,

aiarui aiarui

AIRBORNE AIRBORNE

MOD MOD

LV LV PCM PCM

ABETS ABETS

SEQUENCER/MOC SEQUENCER/MOC

ORDNANCE ORDNANCE

COMMAND COMMAND

nmnii nmnii

AmETiATifi AmETiATifi

ERECTOR

SHUTDOWN SHUTDOWN

PBQPaLAVt PBQPaLAVt

INTERFACE INTERFACE

PRQTCLUUfr PRQTCLUUfr GAIN

KESTRAM KESTRAM

LANDLDK LANDLDK

COMMAND COMMAND

AIRBORNE/GROUND AIRBORNE/GROUND

COMMQMICATTONB COMMQMICATTONB

ma? ma?

RANUC RANUC

STATUS STATUS

PlbOFULSIOH PlbOFULSIOH

PAD PAD

GUIDANCE GUIDANCE

ELECTRICAL ELECTRICAL

FUGIS

SPACECRAFT SPACECRAFT

COORDCMI COORDCMI

LVSS LVSS

MECHANICAL MECHANICAL

O Cn Cn TI8T UJ TITLE

JOINT GUIDANCE AND CONTROL TEST

APPROVAL DATE PREPARED BY: PROJECT ENGINEERING: SYSTEM LEAD: TEST CONDUCTOR: QUALITY* 13 SAFETY: / PILOT SAFETY: CONFIGURATION MANAGEMENT 6555 TH ATW COORDINATION

ENGINEERING PUBLICATIONS: REVISIONS SYM. DESCRIPTION DATE APP*D D Incorporates PCNs 9 thru 15 Ik June 65 Reprinted on II October with PCNs 16 thru 21 inserted,

MODEL (RUN NO. GLV I 42 4

351