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SSD93D0505
STRATEGIC AVIONICS TECHNOLOGY DEFINITION STUDIES
SUBTASK 3-1A, ELECTRICAL ACTUATION (ELA) SYSTEMS
FINAL REPORT
SEPTEMBER 30, 1993
CONTRACT NASg-18880
SOW TASKS A1 - A2
_ Rockwell InternationalSpace Systems Division
STRATEGIC AVIONICS TECHNOLOGY DEFINITION STUDIESSUBTASK 3-1A, ELECTRICAL ACTUATION (ELA) SYSTEMS
FINAL REPORT, SEPTEMBER 30, 1993
PREPARED BY:
C. L. PONDFLIGHT CONTROL SYSTEMS
W. A. McDERMOTTFLIGHT CONTROL SYSTEM
B.T.F. LUMFLIGHT CONTROL SYSTEMS
APPROVED BY:
j. _(_GBURN_/UNCTIONAL MANAGERFLIGHT CONTROL SYSTEMS
Rockwell InternationalSpace Systems Division
Date:
TO:
Internal Letter
September 28, 1993
(Name, Organization, Internal Addresat
K. L. Cureton
SSD-EngineeringD/385-700, 841-AA81
Subject:
#4L_ Rockwell International
No:. 292-600-JO-93-096
FROM: (Name, Orgamzation, lnternal Address, Phone)
J. OgburnSSD-EngineeringD/292-600, 841-FC952-5086
Final Report on Electrical Actuation Study (SWA 29070-45073-
221000,22000)
The purpose of this letter is to present in the enclosure the subject report as
required by the NASA statement of work.
JiJF.lig_hhb(_r_tFo_ional Manager,stems
Avionic Systems
Prepared By: C. L. Pond ,/_.Y/,z_,=/
cc: J. A. McKinney__zB. T. Lum IJW. A. McDermott
PURPOSE OF REPORT
THIS FINAL REPORT PRESENTS THE RESULTS OF AN ELECTRICAL ACTUATION (ELA) SYSTEM
STUDY (SUBTASK TA3-1A) TO SUPPORT THE NASA STRATEGIC AVIONICS TECHNOLOGY DEFINITION
STUDIES.
_L_ Rockwell InternationalSpace Systems Division
CONTENTS
1.0 INTRODUCTION
2.0 ELA TECHNOLOGY DEMONSTRATION TESTING
3.0 ELA SYSTEM BASEUNE
4.0 POWER AND ENERGY REQUIREMENTS FOR SHUTrLE EFFECTOR SYSTEMS
5.0 POWER EFFICIENCY ANO LOSSES OF ELA EFFECTOR SYSTEMS
6.0 POWER AND ENERGY REQUIREMENTS FOR ELA POWER SOURCES
7.0 CONCLUSIONS ANO RECOMMENOATIONS
_WUII IIILtn IlaLIUI Ion
Spece Symnm I)lvlsion
2
1.0 INTRODUCTION
_,_ Roc:kwell |._;_-,,--i_._._Space Systems Olvlslon
3
INTRODUCTION
• ROCKWELL ELA TASKS
• OBJECTIVES
• RESPONSIBILITY
• STATUS ANO ACCOMPLISHMENTS
• SUMMARY OF RESULTS
_ Rockwell Interrmtionalspacesystem revision
4
TA3-1.
ROCKWELL 1993 ELA TASKS(NASA STATEMENT OF WORK)
ELECTRICAL ACTUATION (ELA) SYSTEMS
TA3-1A. ELA SYSTEM TRADES AND BASELINE SELECTION. PERFORM THE FOLLOWINGTRADE STUDIES AND BASELINE SELECTION TASKS:
A1. ELA TECHNOLOGY DEMONSTRATION TESTING
CONTINUE TO SUPPORT THE ELECTRICAL ACTUATION (ELA) DEMONSTRATIONTESTING AT NASA- SELECTED FACILITIES. THE CONTRACTOR WILL: (A)DEVELOP AND COORDINATE TEST PLAN, (B) PERFORM INDEPENDENT REVIEWAND ASSESSMENT OF TEST DATA, AND (C) DOCUMENT AND COOROINATE THETEST RESULTS WITH THE PARTICIPATING NASA ORGANIZATIONS SUCH ASMSFC AND LERC.
A2. ELA SYSTEM POWER EFFICIENCY AND REQUIREMENTS
PERFORM SYSTEMS ENGINEERING DERNITION OF ELA POWER SOURCELOSSES AND REQUIREMENTS FOR SHUTTLE/ORBITER AND SHUTTLE-OERIVEDVEHICLES. THE CONTRACTOR WILL: (A) DEFINE AN ELA BASELINE USINGTHE SPACE SHU'I'rLE AS AN EXAMPLE VEHICLE, (B)OETERMINE POWER ANOENERGY REQUIREMENTS FOR THE SHUTTLE EFFECTOR SYSTEMS, (C).DETERMINE POWER EFFICIENCY AND LOSSES OF THE ELA SYSTEM EQUIPMENT
(POWER SOURCE, ACTUATORS AND CONTROLLERS), AND (D) DOCUMENT THESTUDY RESULTS.
Rockwell Intermtiondspacesystem
5
OBJECTIVES
TASK A1. ELA TECHNOLOGY DEMONSTRATION TESTING
• CONFIRM ADEQUATE STABILITY AND PERFORMANCE OF ELA SYSTEM FOR AEROSPACEAPPLICATIONS.
• OBTAIN TEST OATA TO UPDATE ELA MATH MODELS FOR FAILURE-DETECTION-ISOLATION
(FDI) AND REDUNDANCY - MANAGEMENT (RM) ANALYSES.
TASK A2. ELA SYSTEM POWER EFFICIENCY AND REQUIREMENTS
• DEFINE AN EXAMPLE BASELINE FOR AEROSPACE ELA SYSTEMS.
• ESTABLISH METItODS FOR DETERMINATION OF ELA SYSTEM POWER EFFICIENCY,
LOSSES AND REQUIREMENTS.
• DETERMINE EXAMPLE POWER REQUIREMENTS, EFFICIENCY ANO LOSSES FOR ELASYSTEMS IN AEROSPACE VEHICLES.
Rookwell InternationdSpace Systems Division
6
RESPONSIBILITY
TASK A1. ELA TECHNOLOGY DEMONSTRATION TESTING
RESPONSIBLE ENGINEER: W. A. McDERMOTT
TASK A2. ELA SYSTEM POWER EFRCIENCY AND REQUIREMENTS
RESPONSIBLE ENGINEERS: B.T.F. LUM AND C. L. POND
_,_ Rockwell I,_;¢-,,--_¢,_,-,_Space System Division
7
STATUS AND ACCOMPLISHMENTS
• TASK A1 ELA TECHNOLOGY DEMONSTRATION TESTING• NO TEST SUPPORT WAS REQUIRED DURING THE REPORT PERIOD.
• TASK A2. ELA POWER EFFICIENCY AND REQUIREMENTS
• REVIEWED THE REQUIREMENTS FOR SPACE SHUTTLE EFFECTOR SYSTEMS.
• PERFORMED PRELIMINARY ELA TRADES AND SELECTION.
• DEFINED AN ELA SYSTEM BASELINE.• COMPILED, REVIEWED AND ANALYZED INPUT DATA FOR DETERMINATION OF POWER.
AND ENERGY REQUIREMENTS FOR ELA EFFECTOR SYSTEMS.• DETERMINED POWER EFFICIENCY, LOSSES AND REQUIREMENTS.
Rockwell InternationalSpace Systems Division
8
SUMMARY OF RESULTS
• SECTION 2 PRESENTS THE STATUS AND RESULTS OF A CONTINUING SUPPORT TO NASA ELATECHNOLOGY DEMONSTRATION TESTING.
• SECTION 3 PRESENTS THE ELA SYSTEM CONCEPTS, COMPONENT TRADES AND BASELINESELECTION, USING THE SPACE SHUTTLE AS AN EXAMPLE VEHICLE.
• SECTION 4 PRESENTS THE METHODS, DATA ANO RESULTS FROM THE DETERMINATION OFPOWER ANO ENERGY REQUIREMENTS FOR THE SHUTTLE EFFECTOR SYSTEMS.
• SECTION 5 PRESENTS THE METHODS, OATA AND RESULTS FROM THE DETERMINATION OF
POWER EFFICIENCY AND LOSSES OF ELA SYSTEM EQUIPMENT (POWER SOURCES, ACTUATORS,CONTROLLERS ANO EFFECTORS).
• SECTION 6 PRESENTS THE METHODS, DATA AND RESULTS FROM THE DETERMINATION OF POWERANO ENERGY REQUIREMENTS OF ELA POWER SOURCES.
• SECTION 7 PRESENTS THE STUDY CONCLUSIONS AND RECOMMENDATIONS.
InternationdSpace Systems Division
9
2.0 ELA TECHNOLOGY DEMONSTRATION TESTING
_ Rockwell InternationalSpace Systems Olvlslon
10
STATUS AND RESULTS OF ELA TECHNOLOGY DEMONSTRATION TESTING
• NO TEST SUPPORT WAS REQUIRED DURING THE REPORT PERIOD.
_,_ Rockwell InternationalSpace Systems Division
11
3.0 ELA SYSTEM BASELINE
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SpeceSystemsOlvlslon
12
ELA BASELINE FOR SHUTTLE EFFECTOR SYSTEMS
• EFFECTOR SYSTEM DESCRIPTION
• ELA SYSTEM CONCEPTS
• ELA TRADES AND SELECTION
• EXAMPLE ELA BASELINE
_,_ Rockwell InternationalSpace Systems Division
13
EFFECTOR SYSTEM DESCRIPTION Page1
• THE FLIGHT-CONTROL EFFECTORS IN THE SPACE SHUTTLE ARE DEPICTED IN CHART 16.THESE EFFECTORS ARE:
• GIMBALLED SOLID ROCKET MOTORS (SRM'S), SPACE SHUTTLE MAIN ENGINES (SSME'S)AND ORBITAL MANEUVERING SYSTEM (OMS) ENGINES FOR THRUST VECTOR CONTROL
(TVC),
. AEROSURFACES (ELEVONS, RUDDER, SPEEDBRAKE AND BODYFLAP) FOR ATMOSPHERICFLIGHT CONTROL, AND
• AUXILIARY CONTROLS WHICH INCLUDE SSME PROPELLANT VALVES, EXTERNAL-TANK
(ET) UMBILICAL RETRACTS, BRAKES, NOSEWHEEL STEERING, AND NOSE AND MAINLANDING GEAR UPLOCKS AND STRUTS.
THE SRM-TVC EFFECTORS ARE IN THE SOLID ROCKET BOOSTERS (SRB'S). THE OTHER
EFFECTORS ARE IN THE ORBITER.
• EXCEPT FOR THE OMS-'rvc, THE SHUTTLE EFFECTORS ARE PRESENTLY DRIVEN BY HYDRAULICS.THE OMS-TVC IS DRIVEN BY ELA. THE PROPOSED CHANGE IS TO REPLACE ALL THE SHUTTLEHYDRAULIC EFFECTOR SYSTEMS WITH ELA. STUDIES AT NASA AND ROCKWELL TO-DATE
INOICATE THAT REPLACEMENT OF AUXILIARY POWER UNITS (APU'S) AND HYDRAUUCS INTHE SHUTTLE ORBITER WITH ELA WOULD IMPROVE SAFETY, RELIABILITY, WEIGHT, OPERATIONALCOST, TURNAROUND TIME AND VEHICLE HEALTH MONITORING.
• A TYPICAL EFFECTOR SYSTEM IS ILLUSTRATED IN CHARTS 17 (PICTORIAL) AND 18 (BLOCKDIAGRAM). IN GENERAL, THE SYSTEM CONSISTS OF AN EFFECTOR, ACTUATOR ASSEMBLY,ELECTRONIC CONTROLLERS AND THE LOCAL MOUNTING STRUCTURES. THE ELA EFFECTORSYSTEM IS DRIVEN BY AN ELECTRICAL POWER SOURCE.
Rodmdsymm OWlm.
14
EFFECTOR SYSTEM DESCRIPTION
• THE FUNCTIONS OF THE EFFECTOR SYSTEMS ARE:
• PERFORM PosmONING CONTROL OF THE SHUTTLE EFFECTORS• PROVIDE FAILURE DETECTION ANO ISOLATION (FDI). THIS INCLUDES
FAILURES OCCURRING WITHIN THE EFFECTOR SYSTEM, ANO FAILURESOCCURRING IN INTERFACING SYSTEMS (DATA PROCESSING, ELECTRICALPOWER, ETC.)
• THE TOP-LEVEL REQUIREMENTS FOR THE DESIGN OF THE SHUTTLEEFFECTOR SYSTEMS ARE DESCRIBED IN CHART 19. THESE REQUIRE.
MENTS INCLUDE THE TOTAL NUMBER OF EFFECTORS (OR ACTUATORS)REQUIRED, THE NUMBER OF REOUNDANT CONTROL CHANNELS PERACTUATOR, AND THE MISSION PHASES WHEN THE EFFECTORS ARE
UTILIZED. THE PERFORMANCE REQUIREMENTS (RATES, LOAOS, POWER,ETC.) ARE DISCUSSED IN SECTIONS 4 THROUGH 6.
Rockwd Imernationdspacesymm oh,mem
15
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ILLUSTRATION OF ELA AND HYDRAULIC EFFECTOR SYSTEMS
HYDRAULIC (P.,RESENT) ELA (PROPOSED)
_dL_ Rockwell InternalimNdSpece 8yslems Division
POWERSOURCE
GEARS
CONTROLH
17
BLOCK DIAGRAM OF A SHUTTLE EFFECTOR SYSTEM
HYDRAULIC (_PRESENT)
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_ RockwellInternational8pece 8yllems Olvlslon
18
TOP-LEVEL REQUIREMENTS FOR SPACE SHUTTLE EFFECTOR SYSTEMS
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• _ BAI_KQ_ MAY.BE USI_D FQI_ _:AI I:1)UPI.QCK
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Rockwell InternaOonalSpace Systems Division
19
ELA SYSTEM CONCEPTS FOR SHUTTLE EFFECTORS
• OVERALL ARCHITECTURE
• SYSTEM CONFIGURATIONS
• SYSTEM INTERFACES
• COMPONENTS REQUIRED
• EQUIPMENT LOCATIONS
_,_ Ro(:kwell Intemationd
20
ELA ARCHITECTURE FOR SHUTTLE EFFECTORS
_mn,
• THE OVERALL ELA ARCHITECTURE FOR THE SHUTTLE EFFECTORS IS DEPICTED IN CHART 22.
• THE ELA ARCHITECTURE DESIGN IS GENERALLY BASED ON THE PROVEN CONCEPTS OF THE
PRESENT SHUTTLE EFFECTOR SYSTEMS.
• THE EXISTING FUEL CELLS IN THE ORBITER PROVIDE THE LOW-VOLTAGE POWER FOR THE
EXISTING AVIONICS AND THE ELA CONTROLLERS.
• THE ADDED POWER SOURCES IN THE ORBITER AND EACH_OFTHE SRB'S PROVIDE THE HIGH-
VOLTAGE (TYPICALLY 270 VDC) POWER FOR THE ELA MOTOFI5
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• THSpOWERCOmO.LSRSCO.COLTHSELSCXR,CALCUR.E_.mOUX.E_)WE.SOUeCSTHROUGH THE MOTOR WINDINGS IN THE ACTUATOR ASSEMI_LY _u"_l.
• THE MOTORS ANO THE GEARING IN THE ACTUATOR ASSEMBLY DRIVE THE EFFECTOR ,'=-OR
POSITIONING CONTROL.
(_ Rodmd Intemaeo.dspacesymm o_oa
21
ARCHITECTURE OF ELA EFFECTOR SYSTEMS
IEXISTING I NEW DEVELOPMENT | EXISTING
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Space Syslem DIvMIon
22
ELA SYSTEM CONFIGURATIONS
• THE REQUIRED ELASYSTEM CONFIGURATIONS FOR THE SHUTTLE EFFECTORS ARE DESCRIBEDIN CHARTS 24 THROUGH 28. THESE SYSTEM CONFIGURATIONS ARE:
• FOUR REDUNDANT CONTROL CHANNELS WITH FAILURE DETECTION AND ISOLATION
(FDI), AND
• TWO REDUNOANT CONTROL CHANNELS IN AN ACTIVE-STANDBY ARRANGEMENT.
• DETAILED DESCRIPTIONS OF THE COMPONENTS AND SIGNAL FLOWS OF EACH ELA SYSTEMCHANNEL ARE PRESENTED IN CHARTS 29 THROUGH 53.
• THE ELA SYSTEM CONFIGURATIONS WERE DESIGNED BASED ON THE REDUNDANCY REQUIRE-MENTS SPECIFIED IN CHART 19.
• THE ELA CONCEPT IS GENERALLY SIMILAR TO THE PRESENT HYDRAULIC ACTUATION WITHONE DISTINCT DIFFERENCE. THE ELA SYSTEM USES ELECTRICALCUR_RE.N.-r._S,IGN.ALS WWHj_-I_ONZ-AV OC UrtlmldA//V IMCIN-7FRO FOR DETERMINATION OF FAILURE, THE HYUHAULIU AWl
U_S_'S_S_E'C_)'ND"A"R_[)I'FFF"R-E-I_I'I-ALPRESSURES WHICH ARE NOMINALLY ZERO.
_l InternationdSpace Systems Division
23
FOUR CHANNEL ELA SYSTEM
ii
• THE CONCEPT OF A FOUR-CHANNEL ELA SYSTEM IS ILLUSTRATED IN CHART 25. THE SYSTEMUSES FOUR REDUNDANT CONTROL CHANNELS TO DRIVE THE ACTUATOR AND HENCE THEEFFECTOR. THE FOUR CHANNEL SYSTEM PROVIDES TWO-FAULT-TOLERANT RELIABILITY
(FAIL OPERATIONAL/FAIL SAFE).
• FOR PosmONING CONTROL, THE POSITION COMMAND, ACTUATOR POSITION AND MOTORRATE ARE SUMMED AND COMPENSATED IN THE SERVO PROCESSOR TO FORM A CU,.1RENTCOMMAND TO DRIVE THE MOTOR VIA THE POWER CONTROLLER. THE GEAR TRAIN SUMSTHE MOTOR TORQUE OUTPUTS FROM THE CHANNELS TO DRIVE THE ACTUATOR. EQUALI-ZATION FEEDBACK MAY BE ADDED IF REQUIRED TO COMPENSATE FOR THE REDUNDANTSYSTEM TOLERANCES.
• FOR FAILURE DETECTION AND ISOLATION, THE FDI PROCESSOR IN EACH CHANNEL USES THEDIFFERENCES IN MOTOR CURRENT COMMANDS RETURNED FROM THE POWER CONTROLLERTO DETERMINE FAILURE. THE FAULT ISOLATION COMMAND LOGIC IS TURNED ON (FAILUREINDICATION) WHEN THE CURRENT FAULT MOTOR DETECTS A FAILED CONDITION. THE ISO-LATION COMMAND ISOLATES THE CHANNEL FROM DRIVING THE ACTUATOR BY ZEROING THECOMMAND IN THE POWER CONTROLLER AND REMOVING POWER TO THE MOTOR.
• THE SEI:IVO AND FI)I CONTROLLER ALSO PROVIDES ACTUATOR PosmoN, MOTOR CURRENTAND OPERATING STATUS OF THE CHANNEL TO THE FLIGHT COMPUTER.
• THE INTERFACES BETWEEN CHANNELS AND BETWEEN THE SERVO & FDI CONTROLLERSAND POWER CONTROLLERS ARE ILLUSTRATED IN CHART 26.
_ Rodcwell InternatimmlSpace Systems Division
24
BLOCK DIAGRAM OF FOUR-CHANNEL ELA SYSTEM
Flight
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_,_ Rockwell IntemetkmelSpace Systems mvi,lon
25
SCHEMATIC DIAGRAM OF FOUR-CHANNEL ELA SYSTEM
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_dLq Rod_ell IntemationadSlX_eSystm Division
_'nl_l_R _mmm.y
26
TWO CHANNEL ELA SYSTEM
• THE CONCEPT OF A TWO-CHANNEL (ACTIVE-STANDBY) ELA SYSTEM IS ILLUSTRATED INCHART 28. EACH CHANNEL OF THE ACTIVE-STANDBY SYSTEM IS ESSENTIALLY THE SAMEAS A CHANNEL OF THE FOUR CHANNEL SYSTEM EXCEPT THAT ACTUATOR POSITION IS USED
FOR FDI INSTEAD OF MOTOR CURRENT.
• THE ACTUATOR AND EFFECTOR ARE DRIVEN BY ONE CHANNEL AT A TIME, NORMALLY THEPRIMARY CHANNEL. IF THE PRIMARY CHANNEL FAILS, IT IS SHUT DOWN AND A SIGNAL IS
TRANSMITTED TO THE STANDBY CHANNEL TURNING IT ON. IF THE PRIMARY CHANNEL IS
RESTORED, THE SIGNAL IS REMOVED FROM THE STANDBY WHICH TURNS IT OFF.
• FOR FAILURE DETECTION AND ISOLATION, THE FDI PROCESSOR IN THE ACTIVE CHANNELCOMPARES THE POSITION WITH THE COMMAND TO DETERMINE FAILURE AND GENERATEA POSITION FAULT SIGNAL. THE POSITION FAULT SIGNAL IS USED THE SAME WAY AS THECURRENT FAULT SIGNAL IS USED IN THE MULTI-CHANNEL SYSTEM TO ISOLATE THE FAILED
CHANNEL.
• THE ACTIVE-STANDBY SYSTEM PROVIDES FAIL-OPERATE CAPABILITY. HOWEVER, THETRANSIENT RESPONSE IN RESPONSE TO THE FAILURE AND SUBSEQUENT FDI ACTION
MAY BE GREATER.
Rockwell IntemalimadSpace Systems Division
27
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GEAR TRAIN AND ACTUATOR)• POWER SOURCE (OR SUPPLY)
W Rockwell Internatiomdspace symm m_on
29
SERVO AND FDI CONTROLLER
THE SERVO AND FDI CONTROLLER (ILLUSTRATED IN CHAITIr 31 ) PROVIDES THE FOLLOWINGELECTRONICS:
DIGITAL SERVO PROCESSOR TO PROCESS POSITION ERROR AND MOTOR CURRENTFEEDBACK TO GENERATE MOTOR CURRENT COMMAND TO POWER CONTROLLER.THE SERVO PROCESSOR IS DESCRIBED IN CHARTS 32 THROUGH 34.
DIGITAL FDI PROCESSOR TO PROCESS POSITION ERROR, MOTOR CURRENT FEEDBACKS,ELA FAULT STATUS AND FDI COMMANDS TO GENERATE ISOLATIONS COMMANDS TOPOWER PROCESSOR. THE FDI PROCESSOR IS DESCRIBED IN CHARTS 35 THROUGH 40.
• DIGITAL INTERFACE UNITS TO REFORMAT DATA FOR TRANSFER BETWEEN THE CONTROLLERAND OTHER UNITS VIA DATA BUSES.
• FLIGHT CONTROL INTERFACE UNIT.• POWER CONTROLLER INTERFACE UNIT.• ISOLATION INTERFACE UNIT.• CROSS-CHANNEL INTERFACE CONTROL UNIT.
• POWER SUPPLY TO PROVIDE POWER IN APPROPRIATE FORM TO CONTROLLERELECTRONICS.
• CIRCUITRY TO DETECT POWER SUPPLY FAULTS AND COMBINE STATUS WITH FAULTSTATUS FROM FDI PROCESSOR.
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• IN THE SERVO PROCESSOR (ILLUSTRATED IN CHART 33), THE PosmoN COMMAND, ACTUATORPosmoN AND ROTOR RATE ARE COMPENSATED (BY FILTERS, LIMITERS, ETC.) AND COMBINEDTO FORM A PRELIMINARY MOTOR CURRENT COMMAND.
MOTOR_URRENT SELECTION AND EQUALIZATION (SEE CHART 34) ARE ADDED TO EACHCHANNEL FOR COMPENSATION OF REDUNDANT-SYSTEM TOLERANCES AFFECTING THECHANNEL EQUALIZATION APPLIES ONLY TO MULTIPLE PARALLEL CHANNELS AND DOESNOT APPLY TO ACTIVE-STANDBY SYSTEMS. IT IS NOT NEEDED WITH PRECISION ENCOOERS.
• MOTOR CURRENT FEEDBACK SIGNALS (FOR TORQUE SUMMING) ARE CROS_STRAPPEDBETWEEN CHANNELS, ANO A SELECTION FILTER SELECTS SECOND LARGEST MAGNITUDE.
THE SELECTED SIGNAL IS SUBTRACTED FROM THE IN-CHANNEL SIGNAL AND THEDIFFERENCE IS PROCESSED THROUGH A DEADBAND, A FILTER AND A UMITER TO FORMAN EQUAUZATION FEEDBACK.
• THE EQUALIZATION FEEDBACK IS SUBTRACTED FROM THE PRELIMINARY CURRENTCOMMAND TO FORM THE FINAL MOTOR CURRENT COMMANO.
• THE FINAL MOTOR CURRENT COMMAND IS SENT TO THE POWER CONTROLLER.
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FDI PROCESSOR
• IN THE FDI PROCESSOR (ILLUSTRATED IN CHART 36), THE SIGNAL D FROM THE FAULTISOLATION COMMAND LOGIC IS TURNED ON (FAILURE INDICATION) WHEN THE PosmON-FAULTSIGNAL A (SEE CHARTS 36 AND 38) IS TURNED ON OR WHEN THE CURRENT-FAULT SIGNAL B(SEE CHARTS 36 AND40) IS TURNED ON. FOR THE TWO CHANNEL ELA, O FOR BACKUP IS ONALSO WHEN THE PRIMARY CHANNEL ISOLATION COMMAND IS OFF (PRIMARY CHANNEL GOOD).
D REMAINS ON UNTIL THE RESET COMMAND IS TURNED ON. THE RESET TURNS D OFF ANDRESETS THE CURRENT FAULT MONITOR. FOR THE TWO CHANNEL ELA, D FOR BACKUP ISTURNED OFF WHEN THE PRIMARY CHANNEL ISOLATION COMMAND IS TURNED ON.
THE OUTPUT ISOLATION COMMAND (G) IS ON IF THE SIGNAL D IS ON, THE ISOLATION COMMANDFROM THE FLIGHT COMPUTER IS ON OR IF THE POWER SUPPLY FAULT MONITOR SIGNAL (F) ISON. OTHERWISE THE OUTPUT ISOLATION COMMAND IS OFF.
• THE SERVO AND FDI CONTROLLER ALSO PROVIDES AN OPERATIONAL STATUS OF THECHANNEL TO THE FLIGHT COMPUTER.
Rockwell Imermlkxtdsymm o mo.
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FDI PROCESSOR
(CHANNEL 1, TYPICAL)
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_I_ P,odonll Interrmtlondspece sy=_ mvadon
36
PosmoN FAULT MONITOR FOR FDI PROCESSOR
(SEE CHART 38 FOR ILLUSTRATION)
ill I
THE PosmoN FAULT IS USED ONLY WITH TWO-CHANNEL SYSTEMS.
• IF PosmoN FEEDBACK AGREES WITH COMMAND WITHIN THRESHOLD LEVEL 8L, THE COUNTERSAND FAULT FLAG (A) ARE RESET.
• IF PosmoN ERROR EXISTS AND RATE IS OPPOSITE OF COMMAND FOR SPECIFIED TIME LD, THEFAULT FLAG (A) IS SET.
• IF POSITION ERROR EXISTS AND RATE IS IN RIGHT DIRECTION BUT IS TOO LOW FOR SPECIFIEDTIME LR, THE FAULT FLAG IS SET.
• PARAMETER VALUES OF POSITION FAULT MONITOR ARE DETERMINED FROM THE FOLLOWINGCRITERIA:
POSITION THRESHOLD 8 L GREATER THAN VARIATIONS DUE TO NORMAL OPERATIONAND TOLERANCES AND LESS THAN MAXIMUM ALLOWABLE POSITION ERROR MINUSTOLERANCES
• RATE THRESHOLD RLGREATER THAN VARIATIONS DUE TO NORMAL OPERATION ANDLESS THAN AVAILABLE RATE MINUS TOLERANCES
• COUNTER INCREMENTS AND COUNT LIMITS LARGE ENOUGH TO PREVENT REACTION TOSPURIOUS NOISE AND SMALL ENOUGH TO PREVENT EXCESSIVE POSITION TRANSIENT
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MOTOR-CURRENT FAULT MONITOR FOR FDI PROCESSOR
(SEE CHART 40 FOR ILLUSTRATION)
• THE MOTOR-CURRENT FAULT MONITOR IS USED ONLY WITH FOUR-CHANNEL SYSTEMS.
• MOTOR CURRENT FEEDBACK IN CHANNEL 1 OF TORQUE-SUMMED EMA USED AS ILLUSTRATION.
• WHEN CURRENT FEEDBACK DISAGREES WITH SELECTED CURRENT FEEDBACK BY MORE THANFAULT THRESHOLD IT, INTEGRATOR COUNTS UP AT RATE OF Nu.
• WHEN CURRENT FEEDBACK AGREES WITH SELECTED CURRENT WITHIN THRESHOLD IT,INTEGRATOR COUNTS DOWN AT RATE OF ND.
• WHEN INTEGRATOR COUNTS UP TO FAULT LIMIT NL, FAULT FLAG B IS SET.
• RESET COMMAND RESETS INTEGRATOR TO ZERO.
• THE PARAMETER VALUES OF MOTOR-CURRENT FAULT MONITOR ARE DETERMINED FROM THE
FOLLOWING CRITERIA:
• FAULT THRESHOLD IT GREATER THAN VARIATIONS DUE TO NORMAL OPERATION ANDTOLERANCES AND L_SS THAN AVAILABLE CURRENT MINUS TOLERANCES
• COUNT UP INCREMENT N u AND CURRENT FAULT LIMIT NL LARGE ENOUGH TO PREVENTREACTION TO SPURIOUS NOISE AND SMALL ENOUGH TO PREVENT EXCESSIVE POSITION
TRANSIENT
• COUNT DOWN INCREMENT ND ADJUSTED TO DETECT OSCILLATIONS BUT NOT NOISE
• QUASI-MIDDLE SELECT LOGIC SELECTS SECOND HIGH MAGNITUDE
_ P,ockwell InlemdonelSpace Systems Division
39
MOTOR-CURRENT FAULT MONITOR FOR FDI PROCESSOR (CHANNEL 1, TYPICAL)
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_i_ Flockwell IntornalionalSpace System Olvlslon
40
POWER CONTROLLER
THE POWER CONTROLLER (ILLUSTRATED IN CHART 42) USES THE CURRENT COMMAND FROMTHE SERVO CONTROLLER TO CONTROL THE CURRENT FLOWING FROM THE POWER SUPPLYTHROUGH THE MOTOR WINDINGS. IN GENERAL, THERE ARE TWO TYPES OF POWER CONTROLLER:
• DIRECT-CURRENT (DC) SWITCH-MODE POWER CONTROLLER
• HIGH-FREQUENCY ALTERNATING-CURRENT (AC) POWER CONTROLLER
THESE CONTROLLERS ARE OESCRIBED IN CHARTS 43 THROUGH 46. IT SHOULD BE NOTED THATEITHER THE DC CONTROLLER OR THE AC CONTROLLER AS DESCRIBED CAN DRIVE SYNCHRONOUS
MOTORS (PERMANENT MAGNET OR RELUCTANCE). EITHER CONTROLLER CAN ALSO DRIVE ANINDUCTION MOTOR WHICH IS NONSYNCHRONOUS. IN THIS CASE, THE SEQUENCING OF POWERFROM PHASE TO PHASE IS NOT REFERRED TO AS COMMUTATION. MOTOR RATE MAY BE USED
INSTEAD OF ROTOR POSITION.
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DC SWITCH-MODE POWER CONTROLLER
(SEE CHART 44 FOR ILLUSTRATION.)i i
• THE ERROR SIGNAL (DIFFERENCE BETWEEN COMMANDED CURRENT AND MOTOR CURRENT)CAUSES THE PULSE WIDTH MODULATOR INCLUOING LOOP STABILIZATION COMPENSATIONAND MOTOR-CONTROL SCHEME SUCH AS FIELD ORIENTED CONTROL TO PRODUCE:
• A SERIES OF ON-OFF PULSES WHOSE ON-WIDTHS ARE PROPORTIONAL TO THEMAGNITUDE OF THE ERROR SIGNAL.
• A DIRECTION COMMAND TO CONTROL DIRECTION OF ACTUATOR SLEW.
• THE ON-OFF PULSES CONTROL A SWITCH FOR VARYING THE CURRENT FLOWING FROM THEDC POWER SUPPLY (270 VDC FOR EXAMPLE) THROUGH THE MOTOR WINDINGS.
• THE COMMUTATION LOGIC USES THE DIRECTION COMMAND AND THE ROTOR PosmoNFEEDBACK TO DETERMINE THE ON-OFF SEQUENCE OF THE POWER SWITCHES.
• THE SEQUENCED POWER SWITCHES CONVEFIT THE CURRENT SUPPLY TO THE APPROPRIATETHREE-PHASE POWER TO DRIVE THE MOTOR.
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(SEE CHART 46 FOR ILLUSTRATION.)
• THE OC VOLTAGE FROM THE POWER SUPPLY IS CONVERTED INTO A HIGH-FREQUENCY (SUCHAS 20 KHZ) AC VOLTAGE (380 VAC FOR EXAMPLE)
• THE ERROR SIGNAL (DIFFERENCE BETWEEN COMMANDED CURRENT AND MOTOR CURRENT)CAUSES THE POPULATION DENSITY MODULATOR (PDM) INCLUDING LOOP STABILIZATIONCOMPENSATION AND MOTOR CONTROL SCHEME SUCH AS FIELD ORIENTED CONTROL TOPRODUCE:
• A SERIES OF ON-OFF PULSES WHICH DRIVE A SWITCH FOR CONTROLLING THE AC CURRENTFLOWING THROUGH THE MOTOR WINDINGS.
• A DIRECTION COMMAND TO CONTROL THE DIRECTION OF ACTUATOR SLEW.
• THE ZERO-CROSSING DETECTOR INSURES THAT POWER SWITCHING OCCURS ONLY AT ZEROCURRENT TO MINIMIZE STRESS ON THE SWITCH.
• THE COMBINED FUNCTION OF THE PDM, ZERO-CROSSING DETECTOR, SWITCH, AND RECTIFIERRESULTS IN A SERIES OF HALF-CYCLE CURRENT PULSES WITH DENSITY (NUMBER OF PULSESPER SECOND) PROPORTIONAL TO THE ERROR SIGNAL.
• THE COMMUTATION LOGIC USES THE DIRECTION COMMAND AND THE ROTOR POSITION TODETERMINE THE ON-OFF SEQUENCE OF THE POWER SWITCHES.
• THE SEQUENCED POWER SWITCHES CONVERT THE CURRENT SUPPLY TO THE APPROPRIATETHREE-PHASE POWER TO DRIVE THE MOTOR.
RodramaIntenmlim sp_symm_
45
HIGH-FREQUENCY AC POWER CONTROLLER (EACH CHANNEL)
IMITCN
ZEROClOSE|N!OBTECTOI o
" _ POPULATION
ACTUATOR rO|lTlOBFEEDBACK
COHHUT4T|ONDIRBq LOGIC
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_dL_ Rockwell InternationalSpoco Systems Olvlslon
46
ACTUATOR ASSEMBLY: MOTORS, GEAR TRAIN AND ACTUATOR
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_i_ Rockwell InlemelimmlSpeee Sy_lemm Division
47
MOTOR (PAGE 1 OF 2)
THE MOTORS IN THE ACTUATOR ASSEMBLY CONVERT ELECTRICAL POWER (CURRENT FROM
THE POWER SUPPLY) INTO MECHANICAL MOTION (TORQUE) TO DRIVE THE ACTUATOR. IN
GENERAL, THERE ARE THREE TYPES OF MOTORS THAT ARE APPLICABLE TO ELA SYSTEMS
(SEE CHART 50 FOR ILLUSTRATION): PERMANENT MAGNET (OR BRUSHLESS DC) MOTOR,INDUCTION MOTOR AND VARIABLE (SWITCHED) RELUCTANCE MOTOR. THESE ELA MOTORS ARE
DESCRIBED BELOW.
• PERMANENT MAGNET (PM) MOTOR
• THE ROTOR CONTAINS PERMANENT MAGNET (SUCH AS SAMARIUM COBALT) AND THESTATOR CONSISTS OF THREE PHASES OF WINDINGS.
• THE ROTOR POSITION IS MEASURED AND FED BACK TO THE POWER CONTROLLER FORCOMMUTATION LOGIC WHICH SWITCHES THE MOTOR POWER TO THE NEXT PHASE IN
SEQUENCE.
• THE POWER CONTROLLER ADJUSTS THE POWER FOR THE REQUIRED MOTOR CURRENT.
• THE CURRENT IN THE STATOR WINDINGS REACTS WITH THE MAGNETIC FIELD OF THEROTOR TO CREATE TORQUE THAT CAUSES THE ROTOR TO ROTATE.
• INDUCTION MOTOR
• A SQUIRREL CAGE ROTOR IS USED WHICH HAS LONGITUDINAL CONDUCTORS SHORTED
TOGETHER AT THE ENDS OF THE ROTOR.
• THE STATOR CONSISTS OF THREE PHASES OF WINDINGS.
Rockwell InternationalSpace Systems Division
48
MOTOR (PAGE 2 OF 2)
iii
. THE ROTOR RATE OR POSITION IS MEASURED AND FED BACK TO THE POWER CONTROLLERFOR THE LOGIC WHICH SWITCHES THE MOTOR POWER SEQUENTIALLY FROM PHASE TO
PHASE AT THE COMMANDED RATE.
• THE STATOR-INOUCED CURRENT IN THE ROTOR CONDUCTORS CREATES A MAGNETIC FIELDWHICH REACTS WITH THE CURRENT FLOW IN THE STATOR WINDINGS TO FORM TORQUE.
• VARIABLE (SWITCHED) RELUCTANCE MOTOR
• THE ROTOR HAS NO WINDINGS OR PERMANENT MAGNET; THE STATOR HAS THREE PHASES
OF WlNOINGS.
• THE ROTOR AND STATOR HAVE UNEQUAL NUMBERS OF SALIENT (OR EXPLICIT) POLES.
• THE ROTOR POSITION IS MEASURED AND FED BACK TO THE POWER CONTROLLER FORCOMMUTATION LOGIC WHICH SWITCHES THE MOTOR POWER TO THE NEXT PHASE IN
SEQUENCE.
• THE POWER CONTROLLER ADJUSTS THE POWER FOR THE REQUIRED MOTOR CURRENT
• THE CURRENT FLOW IN THE STATOR WINDINGS CREATES A MAGNETIC FIELD WHICH ALIGNSTHE ROTOR POLES WITH THE MAGNETIZED STATOR POLES RESULTING IN TORQUE ANO
ROTOR ROTATION.
l:lockwell Interml Spece Systems Division
49
ELA MOTOR CONCEPTS
A " A
PERMANI_NT MAGNET INOU(_TION
A
o°
VARIAIN_ RELUOTANCE(SWITCHED RELUOIANCEI
B
_dL_ Rodewell IntematiomlSpace Systems Dlvhdon
5O
GEAR TRAIN AND ACTUATOR (SEE CHARTS 47 AND 52 FOR ILLUSTRATION)
THE GEAR TRAIN SUMS THE PARALLEL CHANNELS TOGETHER INTO ONE OUTPUT, ANDCONVERTS MOTOR TORQUE AND ROTATION INTO ACTUATOR FORCE AND DISPLACEMENT.
THERE ARE FOUR TYPES OF CHANNEL-SUMMING ARRANGEMENT: LINEAR TORQUE SUMMED,ROTARY TORQUE SUMMEO, LINEAR VELOCITY SUMMED ANO ROTARY VELOCITY SUMMED.THESE SUMMING ARRANGEMENTS ARE DESCRIBED BELOW.
• TORQUE SUMMING
• ALL MOTORS GEARED DIRECTLY TO SINGLE SUMMING GEAR• OUTPUT TORQUE IS SUM OF INPUT TORQUES - ALL MOTORS RUN AT SAME RATE• SEPARATE MOTORS FOR MECHANICAL SUMMING - EACH MOTOR MAY HAVE CLUTCH
• MULTIPLE SETS OF WINDINGS IN ONE MOTOR FOR MAGNETIC SUMMING, NO CLUTCHES• FAILED MOTOR (OR WINDING) ISOLATED BY REMOVING POWER ANO DISENGAGING CLUTCH,
IF ANY• VELOCITY SUMMING
• COMBINES MOTOR RATES THROUGH DIFFERENTIALS
• OUTPUT VELOCITY IS SUM OF MOTOR RATES, ALL MOTORS PROVIDE SAME TORQUE• EACH MOTOR REQUIRES BRAKE TO PREVENT BACKDRIVING• FAILED MOTOR ISOLATED BY REMOVING POWER AND APPLYING BRAKE
THE ACTUATOR PROVIDES NUT AND SCREW MECHANISM WITH ROLLERS OR BALLS FOR LINEARACTUATORS, AND ROTARY GEARING TO DRIVE HINGE-LINE FOR ROTARY ACTUATORS o
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POWER SOURCE
THE POWER SOURCE PROVIDES HIGH ELECTRICAL POWER (SUCH AS 270 VDC) FOR ELA SYSTEMS,AND LOW ELECTRICAL POWER (SUCH AS 28 VDC) FOR THE SERVO AND FDI CONTROLLERS ANDOTHER AVIONICS IN THE VEHICLE. THE APPLICABLE POWER SOURCE FOR EL& SYSTEMS ARE:PRIMARY (NON-CHARGEABLE) AND SECONDARY (RE-CHARGEABLE) BATTERIES, FUEL CELLSAND AUXILIARY POWER UNIT (APU) AND OTHER POWER GENERATORS.
Space Systems Division
53
ELA SYSTEM INTERFACES
• THE ELA SYSTEMS INTERFACE WITH THE FOLLOWING SHUTTLE VEHICLE SYSTEMS: DATA
PROCESSING, ELECTRICAL POWER DISTRIBUTION AND CONTROL, THERMAL CONTROL,STRUCTURE, DISPLAY ANO CONTROL, AND EFFECTORS. THE INTERFACE REQUIREMENTS AREDESCRIBED BELOW.
• DATA PROCESSING SYSTEM• INTERFACES WITH COMPUTER THROUGH DIGITAL DATA BUSES; MOM'S ARE NOT
REQUIRED• SOFTWARE IN COMPUTER FOR ELA CONTROL ANO DISPLAY
• ELECTRICAL POWER DISTRIBUTION AND CONTROL SYSTEM (ORBITER AND EACH SRB)• CABLES, SWITCHING CONTROL AND INSTRUMENTATION FOR ELA EQUIPMENT• 270 VDC POWER SOURCE FOR MOTOR POWER AND 28 VDC FUEL CELLS FOR LOGIC
POWER
• THERMAL CONTROL SYSTEM (ORBITER AND EACH SRB)• COLD PLATES FOR ELA CONTROLLERS• PASSIVE COOLING FOR ELA ACTUATOR ASSEMBLIES ANO POWER SOURCE
• STRUCTURES (ORBITER AND EACH SRB)• MOUNTING FIXTURES FOR ELA COMPONENTS AND CABUNG
• DISPLAY AND CONTROL SYSTEM• SAME AS FOR EXISTING HYORAULIC SYSTEMS
• EFFECTORS• SAME AS FOR EXISTING HYORAULIC SYSTEMS
Rockwd Intematkx Systm DiVision
54
ELA COMPONENTS REQUIRED FOR SHUTrLE EFFECTOR SYSTEMS
EACH OF
• SERVO AND FDI CONTROLLER 0* 4
• POWER CONTROLLER 2 45
• ACTUATOR ASSEMBLY 2 45
• POWER SOURCE 4 4
• SHAREO WITH ORBITER EFFECTOR SYSTEMS
(_ RockwegIntemationdspecesymm _
55
ELA EQUIPMENT LOCATIONS
CRITERIA FOR DETERMINATION OF ELA LOCATIONS
• ACCEPTABLE RANGE OF VEHICLE CENTER-OF-GRAVITY (CG) FOR FLIGHT CONTROL• AVAILABLE SPACE• ACCESSIBLE FOR SERVICE AND MAINTENANCE• MINIMUM REQUIREMENTS FOR ELA THERMAL CONTROL• MINIMUM EFFECT DUE TO ANY ELECTROMAGNETIC INTERFERENCE (EMI)• MINIMUM CABLING
POTENTIAL ELA LOCATIONS IN ORBITER• SERVO AND FDI CONTROLLERS: SIMILAR LOCATIONS FOR THE PRESENT AEROSURFACE
SERVO AMPLIFIERS (ASA'S) AND ASCENT TVC (ATVC) DRIVERS (SEE CHART 57)• ACTUATOR ASSEMBLIES: SIMILAR LOCATIONS FOR THE PRESENT HYDRAULIC
ACTUATORS• POWER SOURCE: SEE CHARTS 57 AND 58.• POWER CONTROLLERS: SEE CHARTS 57 AND 59.• El_a, CABLING: SEE CHART 60.
• POTENTIAL ELA LOCATIONS IN EACH SRB('rBD)
_ Rockwell ImematiomdSpace Systems I)lv_on
56
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ELA BATTERIES (4-)
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_ Rockwdl IntemalkxadSmee Systemsl:_Islen
58
POTENTIAL EtA LOCATIONS IN ORBITER: VIEW 3, POWER CONTROLLER (PC)
BAY 7 (NEW)
_i_ Rockwell InternationalSpace Systems Division
59
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ELA COMPONENT TRADES AND SELECTION
• POWER SOURCE
• POWER CONTROLLER
• ACTUATOR ASSEMBLY
• SERVO-FDI CONTROLLER
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61
POWER SOURCE TRADES
• PRIMARY (NON-RECHARGEABLE) SILVER-ZINC (AG/ZN) BA1 I'ERIES• HIGH POWER DENSITY (ESTIMATED POWER AND ENERGY DENSITIES: 1100 W/KG & 150WH/KG)• REQUIRES BATTERY REPLACEMENT EACH FLIGHT
• PRIMARY LITHIUM THIONYL CHLORIDE (LI/SOC12 _ BATTERIES• HIGH ENERGY DENSITY (ESTIMATED POWER AND ENERGY DENSITIES: 450 WlKG & 400 WHIKG)
SAFETY CONCERN DUE TO LOW MELTING POINT OF LITHIUMREQUIRES BATTERY REPLACEMENT BUT MAY HAVE ENOUGH ENERGY FOR MORE THAN
ONE FLIGHT
• SECONDARY (RECHARGEABLE) SILVER ZINC BATTERIES• HIGH POWER DENSITY (ESTIMATED POWER AND ENERGY DENSITIES:• BATTERIES RECHARGED RATHER THAN REPLACED FOR FLIGHT
900 W/KG & 100 WH/KG)
• SECONDARY BIPOLAR LEAD ACID BATTERIES• HIGH POWER DENSITY (ESTIMATED POWER AND ENERGY DENSITIES: 1500 W/KG & 57 WH/KG)
• LOW ENERGY DENSITY• HIGH DEVELOPMENT RISK
• ADVANCED FUEL CELLS• LOW WEIGHT IF USED TO POWER EMA'S AND AVIONICS (ELIMINATING PRESENT
FUEL CELLS)• MORE GROUND SERVICING REQUIRED THAN FOR BATTERIES
AUXILIARY POWER UNITS• SAFETY CONCERN DUE TOXIC PROPELLANTS• HIGH GROUND SERVICING REQUIRED
• BASELINE SELECTION: SECONDARY SILVER ZINC BATTERIES
Rockwell InternationalSpace Systems Division
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POWER CONTROLLER TRADES
• DIRECTCURRENT (DC) SWITCH-MODE POWER CONTROLLER
• POWER SWITCHING AT HIGH CURRENT INCREASING STRESS ON TRANSISTORS• LOWER EFFICIENCY AND GREATER COOLING REQUIREMENTS DUE TO SWITCHING
LOSS AT HIGH CURRENT
• HIGH FREQUENCY ALTERNATING CURRENT (AC) POWER CONTROLLER
• POWER SWITCHING AT ZERO CURRENT REDUCING STRESS_0N_TR_ANS.!_ST_ORS^.N.^R S• THEORETICALLY REDUCED LOSSES AND THERMAL REQUIREMENTS QN mHAN_m_' U• THEORETICALLY PROVIDES IMPROVED PERFORMANCE DUE TO LESS STRINGENT
REQUIREMENTS ON SWITCHING TRANSISTORS• REQUIRES DC TO AC POWER CONVERTOR
• BASELINE SELECTION: HIGH FREQUENCY AC POWER CONTROLLER
Rockwd In lWSondspaces_m raw, on
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ACTUATOR ASSEMBLY TRADES - EMA VERSUS EHA
• ELECTROMECHANICAL ACTUATOR (EMA)
MOTORS DRIVE GEAR MECHANISM WHICH DRIVES EFFECTORNATURAL SUMMING METHOD IS TORQUE SUMMING. VELOCITY SUMMING REQUIRESADDmONAL GEARING MECHANISM
• ELECTROHYDROSTATIC ACTUATOR (EHA)
• MOTORS DRIVE REVERSIBLE HYDRAULIC PUMPS TO DRIVE HYDRAULIC PISTONCONNECTED TO EFFECTOR
• ELECTROHYDROSTATIC MECHANISM READILY PERMITS DISENGAGING CHANNEL
AND DAMPING SUBSEQUENT MOTION• PRESENCE OF HYDRAULIC OIL IN ACTUATOR RAISES CONCERN OF POSSIBLE
HYDRAULIC LEAKS• NATURAL SUMMING METHOO IS VELOCITY SUMMING. TORQUE OR FORCE SUMMING
REQUIRES ADDITIONAL HYDRAULIC MECHANISM.
• BASELINE SELECTION: EMA FOR ALL EXCEPT NOSE WHEEL STEERING.EHA FOR NOSE WHEEL STEERING TO OBTAIN OAMPING MODE
_llVlmll IIIIBl_ Ill_ U
Sgxm _ OMa_l_
.S ¸ •
ACTUATOR ASSEMBLY TRADES - MOTOR TYPE
• PERMANENT MAGNET (BRUSHLESS DC) MOTOR
• SERVO CONTROL TECHNOLOGY USING PERMANENT MAGNET MOTOR HAS BEEN DEVELOPED
ANO DEMONSTRATED• SHORTED TURN MAY CAUSE SEVERE HEATING AND LOAD ON MOTOR DUE TO B _.CK EMF
GENERATED BY ROTOR MAGNET
• INDUCTION MOTOR
• COMMERCIAL PRODUCTION WELL DEVELOPEO• SENSITIVITY TO HIGH TEMPERATURE DUE TO HEATING IN ROTOR
VARIABLE RELUCTANCE MOTOR
• NO SHORTED-TURN PROBLEM ONCE MOTOR POWER IS REMOVEO SINCE ROTOR HAS
NO MAGNETS• RUGGED ROTOR CONSTRUCTION (NO MAGNETS OR WlNOINGS ON ROTOR)
• BASELINE SELECTION: VARIABLE RELUCTANCE MOTOR
specesymm
65
ACTUATOR ASSEMBLY TRADES - CHANNEL SUMMING
• MECHANICAL TORQUE SUMMING (4 SEPARATE MOTORS)
• LOW MECHANICAL COMPLEXITY• ALLOWS USE OF CLUTCH TO ISOLATE FAILED MOTOR• POTENTIAL STRUCTURAL LOADING TWICE NOMINAL WITH FAIL OPERATE/FAIL SAFE
EFFECTOR SYSTEM
• MAGNETIC TORQUE SUMMING (4-1N-1 MOTOR)
• SHORTED-TURN PROBLEM IS NOT CORRECTABLE• POTENTIAL STRUCTURAL LOAO TWICE NOMINAL WITH FAIL OPERATE/FAIL SAFE
EFFECTOR SYSTEM• CLUTCHES OR BRAKES CANNOT BE USED• SIMPLE SUMMING MECHANISM
• VELOCITY SUMMING
• TOLERANT OF SHORTED TURNS (DEVIATION OF MOTOR VELOCITY ENGAGES BRAKE)• COMPLEX SUMMING MECHANISM DUE TO DIFFERENTIAL AND BRAKES• REQUIRES BRAKES TO ISOLATE FAILED MOTORS
• BASELINE SELECTION: MECHANICAL TORQUE SUMMING
Rodzweg Interedm speeesymm
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SERVO & FDI CONTROLLER TRADES - DIGITAL VERSUS ANALOG
DIGITAL CONTROL PROCESSING
• DIRECT CONNECTION TO GENERAL PURPOSE COMPUTER (GPC) DATA BUS. MDM'SNOT REQUIRED.
• MORE PROCESSING CAPABILITIES AND EASIER TO CHANGE
• MORE SENSmVE TO ELECTROMAGNET INTERFERENCE (EMI)• REQUIRES ANALOG-DIGITAL (A/D AND D/A) CONVERSIONS
ANALOG CONTROL PROCESSING
• LESS SENSITIVE TO EMI
• NO A/D OR D/A CONVERSIONS REQUIRED EXCEPT FOR GPC INTERFACES• REQUIRES MDM'S TO INTERFACE WITH GPC'S
• LIMITED PROCESSING CAPABILITY AND INCONVENIENT TO CHANGE• REQUIRES SEPARATE CIRCUIT FOR EACH ACTUATOR
BASELINE SELECTION: DIGITAL PROCESSING
( A Roekwe u. .... .. .67
SERVO AND FDI CONTROLLER TRADES . FDI LOCATION
n
. LOCAL FDI (FDI IN SEI:IYO AND FDI CONTROLLER)
• MINIMIZES EFFECTOR SYSTEM FAILURE TRANSIENTS• MINIMIZES GPC SOFTWARE REQUIREMENTS
• ADDS HARDWARE REQUIREMENTS TO SERVO AND FDI CONTROLLERt
• REMOTE FDI (FDI IN GENERAL PURPOSE COMPUTERS)
• REDUCES HARDWARE REQUIREMENTS ON SERVO AND FDI CONTROLLER• INCREASES EFFECTOR SYSTEM FAILURE TRANSIENTS
• INCREASES SOFTWARE PROCESSING REQUIREMENTS ON GPC'S
• BASELINE SELECTION: LOCAL FDI
Rockwell Imemationdsp symm
68
i _::i"_
ELA BASELINE FOR SPACE SHUTTLE EFFECTOR SYSTEMS
(PRELIMINARY)
ELA EFFECTOR SYSTEMS
• REDUNDANCY LEVEL• SERVO AND FDI CONTROLLER• POWER CONTROLLER• ACTUATOR ASSEMBLY• ELA POWER SOURCE• ELA POWER DISTRIBUTION
ELA INTERFACE SYSTEMS
• DATA PROCESSING• DISPLAY AND CONTROL (PANELS)
• STRUCTURE• THERMAL CONTROL
_ Rockwell InternationalSpace Systems Division
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BASELINE FOR ELA EFFECTOR SYSTEM: REDUNDANCY LEVEL
(PRELIMINARY)
• FOUR PARALLEL CHANNELS FOR:
4 SRB TVC, 6 SSME "I'VC, 4 ELEVONS,RUDDER, SPEEDBRAKE, BODYFLAP,4 BRAKES AND 6 ET (FUEL AND OXIDIZER)UMBILICAL RETRACTS
• TWO CHANNELS (PRIMARY/STANDBY) FOR:
.. 40MS TVC
• NOSEWHEEL STEERING
• 15 SSME PROPELLANT VALVES• 3 NOSE- AND MAIN-GEAR UPLOCKS• 3 NOSE- AND MAIN-GEAR STRUTS
(RETRACTS)
• FOIFS PLUS MINIMIZING EFFECTORFAILURE TRANSIENTS (EFFECTORPOSITION EXCURSION DUE TO FAILUREAND THE SUBSEQUENT FDI ACTION)
• FS (SINGLE OMS ENGINE OPERATIONACCEPTABLE)
• FS PLUS DIFFERENTIAL BRAKINGBACKUP
• FS PLUS PNEUMATIC BACKUP• FS PLUS PYRO BACKUP• FS (GROUND OPERATION ONLY)
_aL_ Rockwell Intemationalspace systems Division
7O
BASELINE FOR ELA EFFECTOR SYSTEMS: SERVO AND FDI CONTROLLER
(PRELIMINARY)
ELA BASELINE
• DIGITAL SERVO CONTROLLER CONNECTED TOGPC DATA BUS
• CROSS.CHANNEL DATA TRANSFER
• ACTUATOR FEEDBACK PROCESSING
• GENERATION OF POSITION ERROR
• FAILURE DETECTION AND ISOLATION (FDI)PROCESSING RESIDES IN SERVO AND FDICONTROLLER (LOCAL FDI)
• BUILT-IN SELF TEST
FOUR CONTROLLERS, ONE FOR EACHCHANNEL. EACH CONTROLLERPROCESSES THAT CHANNEL FORALL ORBITER AND SRB ACTUATORS
CONTROLLERS POWERED BY EXISTINGORBITER POWER SYSTEM
RATIONALE/REMARKS
• ELIMINATE THE NEED FOR MULTI-PLEXER -DEMULTIPLEXER (MDM); MORE PROCESSINGCAPABILITIES AND MORE READILY CHANGED
• NEEDED FOR FDI PROCESSING
• NEEDED FOR SERVO AND FDI PROCESSING
• NEEDED FOR MOTOR CONTROL
• MINIMIZE FAILURE TRANSIENTS PLUSELIMINATE REQUIREMENTS FOR FDISOFTWARE IN THE GENERAL PURPOSECOMPUTERS (GPC'S) AND DATA BUS
• FAULT DETECTION OF SERVO AND FDICONTROLLER
• SAVE WEIGHT, SPACE AND CABLES
• EXISTING ORBITER MATCHES VOLTAGEREQUIREMENTS AND IS READILY AVAILABLE
_ Rockwell Internationalspace systems Division
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BASELINE FOR ELA EFFECTOR SYSTEMS: POWER CONTROLLER
(PRELIMINARY)
ELA BASELINE
• ONE CONTROLLER ASSEMBLY FOREACH ACTUATOR ASSEMBLY WITHALL CONTROL CHANNELS FOR THAT
ACTUATOR
• HIGH FREQUENCY (20 KHZ) AC POWER
INPUT
• USES CURRENT COMMAND FROM SERVOAND FDI CONTROLLERS TO CONTROLCURRENT FLOW FROM POWER SOURCE
THROUGH MOTOR WINDINGS
• RECEIVES FDI COMMANDS FROMSERVO AND FDI CONTROLLERS TOISOLATE FAILED CHANNELS
• INTERFACES WITH ELA TRANSDUCERSTO OBTAIN MOTOR ANO ACTUATORFEEDBACK FOR SERVO AND FDI CONTROLLER
• CROSS-CHANNEL CURRENT DATA TRANSFER
RATIONAt F/REMARKS
• SAVE WEIGHT, SPACE AND CABLES
ALLOW ZERO POWER SWITCHING TOREDUCE LOSSES AND TRANSISTOR
STRESS
DRIVE ACTUATOR MOTORS AS REQUIRED
FOR SERVO CONTROL
• REMOVE EFFECTS OF SYSTEM FAILURES
• REQUIRED FOR SERVO CONTROL
• MINIMIZE IMPACT OF WASTED POWERDUE TO FAILED CHANNEL
_ Rockwell Internationalspace system Division
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BASELINE FOR ELA EFFECTOR SYSTEMS: ACTUATOR ASSEMBLY
(PRELIMINARY)
ELA BASELINE
• VARIABLE RELUCTANCE MOTOR
• MOTOR POSITION TRANSDUCERS
• ACTUATOR POSITION DIGITAL ENCODERS
RATIONALE/REMARKS
• SIMPLE ROTOR, NO SHORTED TURNPROBLEM WITH POWER REMOVED
PROVIDE COMMUTATION SIGNALS TOPOWER CONTROLLER FOR CONTROLLINGSWITCHING TRANSISTORS
PROVIDE ACCURATE POSITION FEED-BACK FOR SERVO CONTROL ANDELIMINATE NEED FOR EQUALIZATIONIN SERVO AND FDI CONTROLLER
ALL EXCEPT NOSEWHEEL STEERING:
• MECHANICAL GEAR TRAIN WITHMECHANICAL TORQUE SUMMING
• LOW COMPLEXITY AND WEIGHT, NOCONCERN OF HYDRAULIC LEAKS
NOSEWHEEL STEERING ELA:
ELECTROHYDROSTATIC ACTUATORMECHANISM WITH FORCE SUMMINGAND DAMPER MODE
PROVIDE DAMPER MODE FOR FAILSAFE OPERATION
_,_ib_ Rockwoli Intematim_Space Systems Division
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BASELINE FOR ELA EFFECTOR SYSTEMS: POWER SOURCE
(PRELIMINARY)
ELA BASELINE
• QUAD REDUNDANT BA'I-I'ERIES (FOURFOR ORBITER AND FOUR FOR EACH SRB
• RE-CHARGABLE SILVER ZINC BATTERIES
• 270 VOLTS DIRECT CURRENT (VDC)
PEAK POWER PER BATTERY: 82 KW FOR
ORBITER; 67 KW FOR SRB
ENERGY PER BATTERY: 5.4 KWH FOR ORBITER;
0.8 KWH FOR SRB
RATIONALE/REMARK_
• FAIL OPERATIONAUFAIL SAFE (FO/FS)
• HIGH POWER DENSITY (STUDIES TO DATE SHOWTHAT BATTERY SIZING FOR SHUTTLE IS DRIVENPRIMARILY BY POWER DENSITY RATHER THAN
ENERGY DENSITY).
• COMMON VOLTAGE FOR ELA'S (ORIGINALLYBASED ON RECTIFIED THREE PHASE 115 VAC)
• PRELIMINARY POWER REQUIREMENTS (TO BE
UPDATED IN FINAL REPORT)
• PRELIMINARY ENERGY REQUIREMENTS (TO BE
UPDATED IN FINAL REPORT)
_aL_ Rockwell InternationalSpace Syslems Division
74
BASELINE FOR ELA EFFECTOR SYSTEMS: POWER DISTRIBUTION SYSTEM
(PRELIMINARY)
ELA BASELINe,
FOUR POWER BUSES FOR ORBITEREFFECTORS AND FOUR POWER BUSES
IN EACH SRB
DC TO 20 KHZ ALTERNATING CURRENT
(AC) POWER CONVERSION
SWITCHING CONTROLS AND CIRCUITINTERRUPTERS
PROVIDE VOLTAGE AND CURRENT
MEASUREMENTS AND STATUSSIGNALS TO DPS AND DISPLAYS
POWER CABLES IN ORBITER LOCATED
GENERALLY THE SPACE AS EXISTINGHYDRAULIC LINES
POWER CABLES IN SRB'S TO BE DETERMINED
RATIONALE/REMARKS
• FO/FS FOR 4-CHANNEL EFFECTOR SYSTEMS
AND FAILSAFE (FS) FOR 2-CHANNEL SYSTEMS
• PROVIDE AC TO POWER CONTROLLER FOR RE-DUCED LOSSES AND THERMAL IMPACTS
ALLOW DISCONNECTING BATTERIES FOR RE-PLACEMENT OR RECHARGING OR ISOLATION OF
SHORTS AND OTHER FAULTS
• PROVIDE POEWR STATUS AND FEEDBACK FOR
FLIGHT CONTROL
° AVAILABLE ROUTING SPACE
75
_ Rockwell InternationalSpace Systems Division
BASELINE FOR ELA INTERFACE SYSTEMS: DATA PROCESSING SYSTEM
(PRELIMINARY)
ELA BASELINE
• CONTROLLER INTERFACE WITH GPC DATA BUS
• EFFECTOR COMMANDS
• ELA FEEDBACK DATA
• UPDATED CRT DISPLAYS
• DISPLAY ELA POSITIONS, CURRENTS,VOLTAGES, STATUS, ETC.
• GENERATE ISOLATE OR RESET COMMANDS
° OTHERS (TO BE DETERMINED)
RATIONALE/REMARKS
• ELIMINATE NEED FOR MDM
PROVIDE EIA DATA TO CREW AND FORDOWNLIST
PROVIDE CREW FDI OVERRIDE CAPABILITY
_L_ Rockwell InternationalSpace Systems Division
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BASELINE FOR ELA INTERFACE SYSTEMS: DISPLAYS AND CONTROL(PRELIMINARY)
ELA BASELINE
• PANEL METERS FOR POWER SYSTEMVOLTAGES AND EFFECTOR POSITIONS
• SWITCHING CONTROLS AND CIRCUITINTERRUPTERS
• OTHERS (TO BE DETERMINED)
RATIONALE/REMARKS
• PROVIDE CREW VISIBILITY OF ELA OPERATION
• PROVIDE CAPABILTIY TO ISOLATE FAULTEDCIRCUITS AND DISCONNECT BATTERIES FOR
RECHARGING
_dL_ Rockwell InternationalSpace Systems Division
77
BASELINE FOR ELA INTERFACE SYSTEMS:
(PRELIMINARY)
STRUCTURE
ELA BASELINE
• NEW ELECTRONICS BAY 7 MOUNTED ON
1307 BULKHEAD
• EXISTING ATTACH POINTS FOR PRESENT
HYDRAULIC ACTUATORS
RATIONALE/REMARKS
• PROVIDE MOUNTING SPACE AND COLDPLATE FOR POWER CONTROLLERS
ELA SYSTEM ACTUATOR ASSEMBLIES TO BEDESIGNED TO FIT SAME SPACE AND ATTACHPOINTS AS PRESENT HYDRAULIC ACTUATORS
• ADDITIONAL MOUNTING PROVISIONS TO BE
DETERMINED
LOCATION OF ELA EQUIPMENT:
SEE CHARTS 56 THROUGH 60 FOR EXAMPLE
#_L_ Rockwell InternationalSpace Syslems Division
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BASELINE FOR ELA INTERFACE SYSTEMS:(PRELIMINARY)
THERMAL CONTROL
ELA BASELINERATIONALE/REMARKS
ACTIVE COOLING (COLD PLATE) FORPOWER CONTROLLERS AND SERVO
AND FDI CONTROLLERS
PASSIVE COOLING (AIR OR HEAT SINK) FORBATTERIES AND ACTUATOR ASSEMBLIES
OTHERS (TO BE DETERMINED)
• REQUIRED BY RELATIVELY HIGH HEAT LOSS,
ELECTRONICS SENSITIVE TO HEAT
• LOW HEAT LOSS AND HEAT SINK TO FR M_4E
CAPABILITY
79
_ Rockwell InternationalSpace Systems Division
08
UOISIAIO stuelsA$ eogd$
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Sn31SX8 H0103.-I-I::1 3"11 InHS I:10.-I SIN::In:II:III'I031:I A_DI:I3N3 QNV 1:13MOd 0"lV
POWER AND ENERGY REQUIREMENTSFOR SHUTrLE EFFECTOR SYSTEMS
• THE MECHANICAL (OR OUTPUT) POWER AND ENERGY REQUIREMENTS FOR EACH INDIVIDUALE EFFECTOR SYSTEM ARE NEEDED IN ORDER TO DETERMINE:
SHUTTL SOURCES SEE SECTION 6.0), AND1 POWER AND ENERGY REQUIREMENTS FOR ELA POWER ......... (_.-.-.-...,.,,n,_
I) THE ELA ACTUATOR SIZING (PEAK POWER CAPABILITIES) i;UH/H," ¢:rr,_., zvno.
UIREMENTS INCLUDE: (1) PEAK POWER AND THE DURATION OF THE PEAK, (2)• THE RE_Q_I....... ..,,,., L,,= ,-,,,,COATIN G DURATION, AND (3) POWER AND ENERGY DEMANDS
AVERA_iE PUW_- _Nu Rn_ vr_,,--AS FUNCTIONS OF MISSION TIME (DUTY CYCLES).
• THE METHODS, DATA AND RESULTS FROM THE CALCULATION OF THE SUBJECT REQUIRE-MENTS ARE PRESENTED IN THIS SECTION AS FOLLOWS.
• METHODS: SEE CHART 82.• DATA: SEE CHARTS 83 THROUGH 127.• RESULTS: SEE CHART 128 THROUGH 136.
ED FOR THE CALCULATION WERE OBTAINED FROM FLIGHTS, TESTS, SIMULATIONSTHE DATA US BECAUSE THEY ARE REPRESENTATIVE OF THEAND ANALYSES. THE DATA WERE SELECTED - L CONDITIONS. THE RESULTSSHUTTLE EFFECTOR SYSTEMS UNDER NOMINAL AND OFF NOMINAWILL BE PRESENTED IN THE FINAL REPORT.
_,L_ Rockwell IntemationalSpace Systems Division
81
METHODS FOR CALCULATION OF EFFECTOR SYSTEMPOWER AND ENERGY REQUIREMENTS
THE MECHANICAL OR OUTPUT POWER (P) AND ENERGY (E) REQUIREMENTS FOR EACH EFFECTORSYSTEM OVER AN OPERATIONAL TIME PERIOD T (SECONDS) IN THE SPACE SHUl-rLE MISSION CAN
BE CALCULATED BY THE FOLLOWING EQUATIONS:
P(t) = 2.64 k L (t) R (t) HORSEPOWER (HP)
E(t) = f: P(t) dt HP- SECONDS (HPS)
WHERE L(t) AND R(t) ARE THE EFFECTOR LOAD (MILLION IN-LBS), AND RATE (DEGREES/SECOND)AS FUNCTIONS OF TIME t (SECONDS). THE PEAK POWER REQUIREMENT OCCURS WHEN THEPRODUCT OF L AND R IS MAXIMUM. THE AVERAGE POWER IS DEFINED AS E/T. THE FACTOR k (k>--l)IS ADDED TO ACCOUNT FOR VARIATIONS IN LOADS AND RATES AS THE RESULTS OF MISSIONMANEUVER VARIATIONS, FLIGHT ENVIRONMENT EXTREMES AND OFF-NOMINAL CONDITIONS.
(k MAY BE STATISTICAL AND TIME DEPENDENT.)
_ Rockwell InternationalSpace Systems Division
82
DATA FOR CALCULATION OF EFFECTOR SYSTEM
POWER AND ENERGY REQUIREMENTS
SRM - TVC
SSME - TVC
OMS - TVC
ELEVONS
RUDDER
SPEEDBRAKE
BODYFLAP
AUXILIARY
CONDITIONS
NOMIINALOFF-NOMINAL (TRAJECTORY VARIATIONS)
NOMINALOFF-NOMINAL (3 DISPERSION & ENGINE OUT)
NOMINAL AND OFF-NOMINAL
NOMINAL AND OFF-NOMINAL
NOMINAL AND OFF-NOMINAL
NOMINAL AND OFF-NOMINAL
NOMINAL AND OFF-NOMINAL
OFF-NOMINALENTRY WITH2 RCS JETSFAILURE
NOMINAL AND OFF-NOMINAL
DATA (LOADS & RATES_) T_rPES __
84-87 S, A, T 1, 388-91 S, A 3
93-106 S_ A, T,F 3
107-112 c S, A 3
127 A 3
113 - 120 S, A 3
121 - 122 S, A 3
123 - 124 S, A 3
125 - 126 S, A 3
127 A 2, 3
NOTES: LOADS AND RATES DATA: SEE CHARTS WITH NUMBERS INDICATED.
DATA TYPES: T = TEST; A = ANALYSIS; S= SIMULATION; F= FLIGHT DATADATA SOURCES: 1 = NASA MSFC; 2 = NASA JSC; 3 = ROCKWELL
_,dL_ Rockwell InternationalSpace Systems Division
83
SRM NOZZLE LOAD ON ROCK OR TILT ACTUATOR VERSUSMISSION TIME AND GIMBAL ANGLE
.... Extrapolstlon
SRB
NOZZLE
LOAD,
HILLIOH
IN-LIIS
$.0
4.$
4.0
3.5
3.0
2.5
2.0
1.5
1.0
0:5
0.0
NOZZLE
GIItBAL
ANGLE,
DEGREES
0
NOTE: AN AERODYNAMIC LOAD(ABOUT 0.17 MILLION IN-LBSMAXIMUM) SHOULD BE ADDEDTO THE NOZZLE LOAD IN ORDERTO OBTAIN THE TOTAL LOAD ONTHE ACTUATOR.
2O 40 60 80 100 120
ASCENT HISSION TINE, SECOHOS
_dL_ Rockwell InternationalSpace Systems Olvlslon
84
ESTIMATED SRM SIDELOADS AT IGNITION AND SHUTDOWN
FIRING TEST RESULTS HAVE INDICATED THAT THE SRM SIDELOADS AT IGNITION AND SHUTDOWNDO NOT RESULT IN SIGNIFICANT LOADING OF THE SRM EFFECTOR SYSTEM AND HAVE NOIMPACT ON POWER OR ENERGY REQUIREMENTS.
ACTUATORLOAD IN
MILLION
IN:LBS
TIME HISTORY OF :lOCK ACTUATOR LOAD DURING IGNITION FOR SRB QM4
2.78
1.85
0.92
0
'0.92
-1.85
-2.78t
SRM IGNITION
¢CMVW_
TIME, SECONDS
%,
_L_ Rockwell InternationalSpace Systems Division
85
PosmoN AND RATES OF RIGHT SRM - TVC ACTUATORS DURINGNOMINAL ASCENT
81_,-56FIRSTSTAGEASCENTSIMULATION. NOMINAL MISSION
SRB11LT RK;HT_ ROCK
2.01.5 I I I I
a 0.5
0.0 I i415-I.0 " . I '-I.5 ' I I : _ I
0 20 40 60 80 100 120 140
tIME. SZC
0o.o
-I.0
L -2.0
-3.00
2.0
_1"_"_:_-.._,,
!iv ' l J,,20 40 60 80 I00 120
tIME. SEC
140
..o
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o 20 40 60 eo ioo ,20 ,40TIME, SEC
4
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TIME, SEC
140
_iL_ Rockwell IntemationalSpace System Division
86
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$,J_ Rockwell IntemationalSpace Systems Division
i i
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SRM NOZZLE GIMBAL RATES VERSUS MISSION TIME FOR LEFT-SRB TVC
ACTUATORS WITH,NO FAILURES (150 SIMULATED STS-1 MISSIONS)
jo
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NO san 145
_,_ Rockwell InternationalSpace Systems Division
91
LOADS ON SSME-TVC EFFECTORS (FLIGHT DATA)
THE LOADS (T) ON THE SSME-TVC EFFECTORS (OR ACTUATORS) ARE CALCULATED BY THEFOLLOWING EQUATION:
T - (L_P), (A). (L) IN-LBS
WHERE Z_P IS THE ACTUATOR PRIMARY PRESSURE ( POUNDS PER SQUARE INCH OR PSID);IS THE ACTUATOR PISTON AREA (SQUARE INCHES); L IS THE ACTUATOR MOMENT ARM (INCHES).
Au_ ,,,T^ ,,_=n =rip TH_ P.A/CIILATION ARE AS FOLLOWS: &P S ARE THE MEASURED PRIMARY=n= u_-_ v_ --- ............ . - IGHTS SEE CHARTS 93
THE SHUTTLE STS 1 THROUGH STS 5 FLcH AN -('UPRESSURES OBTAINED. FROM 20.03 FOR OTHER. • TO 24 83 FOR SSME 1 (UPPER ENGINE) PITTHROUGH 97), A IS EQUAL
SSME-TVC ACTUATORS; L IS EQUAL TO 29.74 FOR EACH ACTUATOR.
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124
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_dL_ Rockwell InternationalSpace Systems Division
126
POWER AND ENERGY DATA FOR OTHER EFFECTORS
EFFECTOR
(_TOTAL NUMBER_
OUTPUT POWER MISSION ON DATAREQUIRED, HP TIME, SEC
OMS - TVC'S (6) 1SSME VALVES (15) 3.3, 1.5, 1.1, 1.1,ET UMBILICALS (6) 1.0MAIN GEAR UPLOCKS (2) 5.0NOSE GEAR UPLOCK (1) 2.0MAIN GEAR STRUTS (2) 5.0NOSE GEAR STRUT (1) 2.0NOSEWHEEL STEERING (1) 3.7BRAKES (4) 9.0
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ESTIMATENASA JSC DATANASA JSC DATANASA JSC DATANASA JSC DATANASA JSC DATANASA JSC DATARATE VS LOADNASA JSC DATA
ENERGY REQUIREMENTS ON ORBIT ARE NEGLIGIBLE.
_dL_ Rockwell InternationalSpace Systems Division
127
LOADS ON SRM-TVC EFFECTORS
THE SIGNIFICANT LOADS ON THE SRM-TVC EFFECTORS (OR ACTUATORS) ARE THE NOZZLERESTORING TORQUE (T) AND THE AERODYNAMIC LOAD. THE NOZZLE RESTORING TORQUE ISPRESENTED IN CHART 84 AS A FUNCTION OF NOZZLE ANGLE (A) AND TIME. THE AERODYNAMICLOAD IS A FUNCTION OF THE DYNAMIC PRESSURE (QBR) AND WAS ESTIMATED FROM ANANALYSIS OF WIND TUNNEL DATA. THE TOTAL LOADS (L) ON THE SRM-TVC EFFECTORS ARECALCULATED BY THE FOLLOWING EXPRESSION:
L = T(A)-314.QBR IN-LBS.
700._ 600
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_aL_ Rockwell InternationalSpace Systems Division
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130
SRB TVC ACTUATOR NOZZLE LOAD AND POWER DURING ASCENT
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#1_ Rockwell InternationalSpace Systems Division
131-1
SRB TVC ACTUATOR NOZZLE LOAD AND POWER DURING ASCENT
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_ Rockwell InternationalSpace Systems Division
131-2
SRB TVC ACTUATOR NOZZLE LOAD AND POWER DURING ASCENT
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#_ Rockwell InternationalSpace Systems Division
132-1
SRB TVC ACTUATOR NOZZLE LOAD AND POWER DURING ASCENT
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#_L_ Rockwell InternationalSpace Systems Division
132-2
ESTIMATED LOADS ON SSME-TVC EFFECTORS (SIMULATION DATA)
THE LOADS ON THE SSME-TVC EFFECTORS (OR ACTUATORS) ARE ASSUMED TO INCLUDETHREE PRINCIPAL COMPONENTS: (1) AERODYNAMIC LOAD, (2) GIMBAL FRICTION AND (3)THRUST OFFSET LOAD.
THE ESTIMATED AERODYNAMIC LOADS ARE CALCULATED BY THE FOLLOWING EXPRESSIONS:
PITCH YAWSSME 1 (611.9 - 54.3 PIPOSN)QBR (138.2 -68.3 YIPOSN) QBRSSME 2 -404.1 QBR (-192.4 - 52.9Y2POSN) QBRSSME 3 -404.1 QBR (+192.4 - 52.9Y3POSN) QBR
THE NUMERICAL VALUES ARE BASED ON AN ANALYSIS OF WIND TUNNEL DATA, AND "QBR"IS THE DYNAMIC PRESSURE PRESENTED IN CHART-- .
THE GIMBAL FRICTION IS REPRESENTED AS 223750.SIGN (GIMBAL RATE) WHERE THENUMERICAL VALUE IS BASED ON THE CURRENT MATH MODEL OF THE HYDRAULIC SSME-TVCEFFECTOR SYSTEM ON THE SPACE SHUTTLE.
THE THRUST OFFSET LOAD IS CALCULATED BY THE FOLLOWING EQUATIONS:THRUST: T = 400000 LBMAXIMUM OFFSET: D = 0.6 INNO. OF ACTUATORS: N = 6LOAD PER ACTUATOR: L = TDIJ'N = 100000 IN-LB.
HOWEVER, 240000 IN-LB (L =TD) WAS USED FOR SSME 1 PITCH. THE LOADS USED ARE:
PITCHSSME 1 +240000 IN-LBSSME 2 +100000 IN-LBSSME 3 -100000 IN-LB
YAW+100000 IN - LB- 100000 IN - LB+100000 IN - LB
#dL_ Rockwell InternationalSpace Systems Division
133
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_00.,5
0.00 20 40 60 80 100 120
TIME, SEC
140
#i_ Rockwell InternationalSpace Systems Division
134-1
SSME TVC ACTUATOR LOAD AND POWER DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, 3 SIGMA DISPERSION
SSME #2 PITCHSSME#2 YAW
__. 200.000
0
'_ -200_00
<- -40o_oo-_ -600,000
0 50 100 150
TIME, SEC
200.000
_- 100,0000
_, -100,000
-200,000-300,000
-40o,ooo0
i i
50 100
TIME, SEC
i
150
5.04.03.0
i 1.02.0
0.00 -I.0
-2.00 140
i i 1 i 1 i
20 40 60 80 100 120
TIME, SEC
3.53.02.5
2.01.51.0
_O0.5
0.00 20 40 60 80 100
TIME, SEC
20 140
#_L_ Rockwell InternationalSpace Systems Division
134-2
SSME TVC ACTUATOR LOAD AND POWER DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, 3 SIGMA DISPERSION
SSME #3 PITCHSSME #3 YAW
z' 200,000
0
_,_ -2oo,ooo.=,I
-400,000
-600.000 "
0 ,50 100 150
TIME, SEC
_, 400,000- 300,000
200.000
O_, 100,0000
-100,000-2oo.00o
0
JL
,50 100
TIME, SEC
I
6.05.04.0
i 3.02.01.00.0 ,i
-I.00 20 40 60 80 100 120 140
TIME, SEC
2.5
2.0
1.51.0
_0.5
0.00 20 40 60 80 100 120
TIME, SEC
140
#4L_ Rockwell InternationalSpace Systems Division
134-3
SSME TVC ACTUATOR LOAD AND POWER DURING ASCENT
STS-53SECOND STAGE ASCENT SIMULATION, NOMINAL MISSION
SSME#1 PITCHSSME#I YAW
_ 500.000 I
6 400.000
_..,q 300.000•-, 200.000
__ l°°'°_o0 100 200 300 400 500 600
TIME, SEC
- 300.000•_ 200.000
_ I00.
-100.000 |-200_[]0 ' '
0 I00 200 300 400 500 600
TIME, SEC
5.04.0
i 2.03.0
1.00.o
i l i +
100 200 300 400
TIME, SEC
-I.00 5OO 6OO
0.3
0.3
_00.2
0.2o.1
O 0.10.0
0
A ,100 200 300 400 500 600
TIME, SEC
#_L_ Rockwell InternationalSpace Systems Division
135-1
SSME TVC ACTUATOR LOAD AND POWER DURING ASCENT
STS-53SECOND STAGE ASCENT SIMULATION, NOMINAL MISSION
SSME#2 PITCHSSME#2 YAW
_ 500,000j
400,000
_=I 300,000
•-.* 200,0001<.'°°_°° l
0 '
0 100 200 300 400 500 600
TIME, SEC
_: 400,000
- 300,000I
200,000I
i.,oo I-I00,000|-200,[]001 ,
0 100 200 300 400 5[]0
TIME, SEC
6OO
3.02.5
i 2,01.51.0
_0 0.50.0
-0.50 100 200 300 4[]0 500
TIME, SEC
i
600
3.0 ¸
2.5
_O2.01.51.0
0 0.5
0.0
0
i
IO0 200 300 400 500
TIME, SEC
i
600
#_L_ Rockwell InternationalSpace Systems Division 135-2
SSME TVC ACTUATOR LOAD AND POWER DURING ASCENT
STS-53SECOND STAGE ASCENT SIMULATION. NOMINAL MISSION
SSME#3 PITCHSSME#3 YAW
z_" 500.000
4OO.OOO
_ =1 300.000.-, 200.000
100
0 100 200 3O0 4OO 5O0 600
TIME, SEC
_. 400.000
- 300.000 t
2oo.oool
-IO0,OO0|-200,000 ' '
0 100 200 300 400 500
TIME, SEC
i
8.0
6.04.0
2.0
_0 0.0
-2.00 6OO
l L I I
100 200 300 400 500
TIME, SEC
0.8_0.6
0.4
o 0.20.0
0 100 200 300 4[]0 500
TIME, SEC
600
#_L_ Rockwell InternationalSpace Systems Division
135-3
AEROSURFACE ACTUATOR LOAD AND POWER OUTPUT DURING ENTRY
SEE CHARTS 113-126 FOR HINGE MOMENT AND OUTPUTPOWER VERSUS MISSION TIME DURING ENTRY AND LANDING.
#_L_ Rockwell InternationalSpace Systems Division
136
5.0 INPUT POWER AND ENERGY REQUIRED BY ELA EFFECTOR SYSTEMS
#,_ Rockwell InternationalSpace Systems Division
137
ESTIMATED POWER EFFICIENCY AND LOSSES OFEQUIPMENT IN SHUTTLE ELA EFFECTOR SYSTEMS
--------'-_ p(_Wl::R EFFICIENCY PERCENT (OR POWER L(_.SR_ KW_ RYRTFM _
EFFECTOR SYSTEM " " " EFFICIENC;"
SRM - TVCSSME - TVCELEVONRUDDERSPEEDBRAKEBODYFLAP
90%90%90%90%90%90%
0.5 X 2.644 X T1.0 X 2.644 X T3.0 X 2.644 X T1.4 X 2.644 X T0.6 X 2.644 X T0.3 X 2.644 X T
95%95%95%95%95%95%
80%80%80%95%95%95%
NANA
80%80%80%
68%68%68%65%65%65%
* POWER CONTROLLER LOSS WHILE HOLDING EFFECTOR POSITION AGAINSTCONSTANT LOAD WHERE T - EFFECTOR LOAD IN IN-LB. THIS LOSS WAS NOTINCLUDED IN THE PLOTS.
** ELEVON LOAD INCLUDES 9000 IN-LB FRICTION PLUS HINGE MOMENT.
#_ Rockwell InternationalSpace Svstems Division
138
ESTIMATED POWER EFFICIENCY AND LOSSES OF EQUIPMENT
IN SHUTTLE ELA EFFECTOR SYSTEMS (CONTINUED)
• SERVO & FDI CONTROLLER POWER INPUT IS ESTIMATED TO BE 50W EACH, OR 200W TOTAL.
• ASSUMED EFFICIENCY OF RECHARGEABLE POWER SOURCE IS 90%. THAT IS, THE
RECHARGEABLE SOURCE RETURNS 90% OF ENERGY INPUT BACK TO THE ELA.
• AND LOSSES ARE ESTIMATED BASED ON SHUTTLE EFFECTOR SYSTEMSEFFICIENClES OR ENERGY AVAILABLEAND SIMILAR EQUIPMENT. THERE IS NO ELA TEST DATA ON POWER
TO CONFIRM OR UPDATE THESE VALUES.
• SHUTTLE EFFECTORS TYPICALLY EXPERIENCE AIDING LOADS (REGENERATIONTHE SPACE LE POWER SOURCES
DS MOTORING MODE). RECHARGEABMODE) AS WELL AS OPPOSING .I-ORE ;...(_,.,.,.,,,. v ,..,c-rHDu_n BY THE ELA. REGENERATIVE
RIES CAN AGGP..I"/I=l_r-_= _..v,.,,,,-,.-GENERALLY BATTE ) NERGY MUST BEMODE. WITH NONCHARGABLE POWER SOURCES, THE REGENERATION E
DISSIPATED AND IS NOT AVAILABLE FOR LATER USE.
• SINCE THE FOUR CHANNELS ARE NOMINALLY IDENTICAL, THE ACTUATOR LOAD AND LOSSES
ARE DIVIDED EVENLY BETWEEN THE ACTIVE CHANNELS. THE LOSS PER CHANNEL IS THERE-
FORE THE TOTAL LOSS DIVIDED BY THE NUMBER OF ACTIVE CHANNELS.
Rockwell InternationalSpace Systems Division
139
CALCULATION OF INPUT POWER AND ENERGY FROM OUTPUT POWER AND EFFICIENCY
• THE POWER INPUT INTO THE ELA WITH OPPOSING LOAD IS CALCULATED BY THE FOLLOWING
EQUATION:
KW - 0.746 HPIEFF1
WHERE KW IS INPUT POWER IN KILOWATTS, HP IS OUTPUT HORSEPOWER AND EFFt IS THEELA SYSTEM EFFICIENCY.
• WITH RECHARGABLE BATTERIES AND AIDING LOAD, THE POWER RETURNED TO THE BATTERIESFOR LATER USE BY REGENERATION IS CALCULATED BY THE FOLLOWING EQUATION:
KW - (0.746 HP) (EFF1) (EFF2)
WHERE EFF2 IS THE BATTERY RECHARGING EFFICIENCY.
• ENERGY SUPPLY IS THE TIME INTEGRAL OF THE SUPPLIED POWER.T
KWH - J00KW dt
NOTE THAT THE REGENERATION POWER IS NEGATIVE SUPPLIED POWER.
#_i,_ Rockwell InternationalSpace Systems Division
140
CALCULATION RESULTS: POWER AND ENERGY REQUIREMENTS FOR SHUTTLE EFFECTORS
EFFECTOR PEAK POWER PER ELA, KW
(DURATION. SECONDS_
SRM - TVCSSME - TVC
UPPER PITCHLOWER PITCH
YAWELEVONS
INBOARDOUTBOARD
RUDDERSPEEDBRAKEBODYFLAPMAIN GEAR UPLOCKSNOSE GEAR UPLOCKNW STEERINGBRAKES
12 (1)
7.0 (1)20 (1)13 (1)
8.8 (1)5.2 (1)1.0 (5)15 (1)1.5 (1)5.3 (5)2.1 (5)5.3 (60)13 (60)
* SAME AS FOR PEAK POWER
AVG. POWER PER ELA, KW
(DURATION, MINUTES)
0.42 (2)1st STAGE, 2nd STAGE0.63 (2), 0.08 (6.4)0.69 (2), 0.16 (6.4)0.31 (2), 0.09 (6.4)
0.21 (18)0.09 (18)NEGLIGIBLE
0.22 (18)0.01 (18)
#,dL_ Rockwell InternationalSpace Systems Division
14:1
SRB TVC ACTUATOR INPUT POWER DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, NOMINAL MISSION
LEFTSRBTILTLEFTSRBROCK
" WITH REGENERATION
_ 2 .......
m
-6 ! 1 * L J
0 20 40 60 80 100 120 140
TIME, SEC
2.5
..-0.5
-1.50
i i
20 40 60
WITH REGENERATION
. . . ; ill, I,, - _._
, t t -4
80 100 120 140
TIME, SEC
8
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2
00
NO REGENERATION
1
20 40 60 80 100 120 140
TIME, SEC
25i I2 NO REGENERATION
1.5
o._ i A,,_lnall_l,iRil, ....... ,,,,,,,.,_.0.. . A .....
0 20 40 60 80 100 120
TIME, SEC
i
140
4L4 Rockwell InternationalSpace Systems Division
142
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STS-56FIRSTSTAGE ASCENT SIMULATION, ENGINE 2 FAILED AT LIFTOFF
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#_ Rockwell InternationalSpace Systems Division
144
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STS-56FIRSTSTAGEASCENT SIMULATION, 3 SIGMA DISPERSION
RIGHT SRBTILT RIGHTSRBROCK
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2
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TIME, SEO
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#_ Rockwell InternationalSpace Systems Dlvlslon
146
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SRB TVC ACTUATOR INPUT ENERGY DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, NOMINAL MISSION
LEFTSRBTILTLEFTSRBROCK
0.004
0.0030,0020.001 WITHREGENERATIO
" o
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-- 0 20 40 60 80 I00 120 140
TIME, SEC
o.oo21 _
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-- " _0 2'0 40 60 80 lO0 120
TIME, SEC
_ 0.008
>_ 0.006 __,F-_"
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i i , i
40 60 80 100 120 140
TIME, SEC
0.0030.0025
0.0020.0015
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0
f
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20 40 60
NO REGENERATION
80 I00 120
TIME, SEC
l
140
i
140
#_L_ Rockwell InternationalSpace Systems Division
148
SRB TVC ACTUATOR INPUT ENERGY DURING ASCENT
STS-56FIRSTSTAGEASCENT SIMULATION, 3 SIGMA DISPERSION
LEFTSRBTILT LEFTSRBROCK
0.012 TOOlto.oo8t
i .O,OO2I ' '
0 20 40
/WITH REGENERATION
I I I I
60 80 loo 120
TIME, SEC
140
0.002 lo.oo_t _ _--_o.oo_t .,.J'_
0. _ WITH REGENE IO
_ -0.OO05 i ......-- 0 20 40 60 80 100 120
TIME, SEC
L
140
_ 0.015
0.01
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NO REGENE,RATIO ,i I
0 20 40 60 80 1O0 120 140
TIME, SEC
0.003 To,oo_I°°°2 t
o.oo15t _'_
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- 0 20 40 60 80 I00 120
TIME, SEC
I
140
#_L_ Rockwell InternationalSpace Systems Division
149
,. .: ; ::
SRB TVC ACTUATOR INPUT ENERGY DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, ENGINE 2 FAILED AT LIFTOFF
LEFTSRBTILT
I0.0050.0040.OO3 __ o.oo2t f_'
w 0.00 WITH REGENERATION
-_'°'®"o 2'o 4o _ _'o ,® ,'_o ,_oTIME, SEC
LEFTSRBROCK
I _ 0.0025
0.002
_ 0.0015
0.001 REGENERATION
S o.ooo5L
_- °o _'o ,_ 6o 8'o ,® ,_oTIME, SEC
0.010.OO8>.
0.006_ 0.004
S 0.0020..
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....,-.--J
NO REGENERATION
i l i J
60 80 100 120 140
TIME, SEC
_ 0.00350.003>_ 0.0025
0.002-._ 0.0015
0.0010.0005z 0
0
S NO REGENERATION
, o i i 4
20 40 60 80 I00 120
TIME, SEC
i
140
140
#i_l Rockwell InternationalSpace Systems Division
150
SRB TVC ACTUATOR INPUT ENERGY DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, NOMINAL MISSION
RIGHT SRBTILTRIGHTSRB ROCK
0.00250.002
0.0015
!O° o__ -0.0005
_REGENERATION
ii i i i i L
0 20 40 60 80 lO0 120 140
TIME, SEC
'O'lj0.006
(_ 0.0050.0040.003 WITH REGENERATION0.002
0.001z -0.001 .....
-- 0 20 40 60 80 100
i
120
TIME, SEC
i
140
_ 0.00350.0030.0025
0,0020.0015
O. '
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_-'-_'_ NO REGENE ATION
y ,,,i,,20 40 60 80 100 120 140
TIME, SEC
0.01 I -'-
0.008 J i -_-'_-_-
-_ 0.004 NO REGENERATION
0.0020 .....
-- 0 20 40 60 80 100 120
TIME, SEC
I
140
#i_ Rockwell InternationalSpace Systems Division
151
SRB TVC ACTUATOR INPUT ENERGY DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, 3 SIGMA DISPERSION
RIGHT SRBROCKRIGHTSRBTILT
:_2520 20 ..-.--,",-_
10 10 WITHREGENERATION
5- TH REGENERATION 5
0 ......0 ' ' _ _ -5 40 60 80 100 120 140Z -5 .... ---- 0 20 40 60 80 100 120 140• TIME, SEC
TIME, $EC
..J--- i '--'--15
• ,ot / .o.,o,.,_,.o. "-,o .o.,.,,,,.,,.o.I _ts_. _ .......' ' ' 1 1 140
0 20 40 60 80 TIME, SECTIME, SEC
#i_ Rockwell InternationalSpace Systems Division
151A
_ 0.004
_ 0.0030.002
_ 0,001
_Z 0
SRB TVC ACTUATOR INPUT ENERGY DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION. ENGINE 2 FAILED AT UFTOFF
RIGHT SRBTILT RIGHT SRBROCK
I_F I I i I i i_ i
0 20 40 60 80 I00 120 140
TIME, SEC
O I0.003
WITHREGENERATION_, 0.002
'.' ooo__j ...._ _0.001 1 ......-- 0 20 40 60 80 100 120
TIME, SEC
4
140
_ 0.CX_0.005
_ 0.0040.0030.002
I_ 0.001a.
z_ o0
I I I I i I I
20 40 60 80 I00 120 140
TIME, SEC
0.01 10.008
_ 0.0040.002
Or , ,
0 20 40
f
NO REGENERATION
I I i
60 80 I00
TIME,SEC
l
120
i
140
_,_ Rockwell InternationalSpace Systems Division
152
SSME TVC ACTUATOR INPUT POWER DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, 3 SIGMA DISPERSION
SSME#1 PITCHSSME#1 YAW
_. 6
_42_
_om
-20
WITH REGENERATION
" 4
I I I I I I
20 40 60 80 I00 120 140
TIME, SEC
2.5
1.51
_. 0.500 20
WITH REGENERATION
40 60 80 I00 120
TIME, SEC
_!Illl ,AI,ILALLIIIL,_ n . . . i . . .
0 20 40 60 80 I00 120 140
TIME, SEC
2.5
2L_1.51
0.5
00 20
NO REGENERATION
40 60 80 I00 120
TIME, .SEC
140
¢
140
#_L_ Rockwell InternationalSpace Systems Dlvlslon
153
SSME TVC ACTUATOR INPUT POWER DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, 3 SIGMA DISPERSION
SSME#2 PITCHSSME#2 YAW
5
" 3
1
_o-I
0
WITH REGENERATION
i
I I I i I
20 40 60 80 I00 120 140
TIME, SEC
0 20 40 60 80 I00 120
TIME, $EC
140
43
l
00
I NO REGENERATION
. , . . , -
20 40 60 80 I00 120 140
TIME, SEC
3.5
- 2.i
iO,o0
__INO REGENERATION
20 40 60 80 I00 120
TIME, SEC
140
_i_ Rockwell InternationalSpace Systems Division
154
0.n
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SSME TVC ACTUATOR INPUT POWER DURING ASCENT
STS-56FIRSTSTAGE ASCENT SIMULATION, 3 SIGMA DISPERSION
SSME #3 PITCHSSME #3 YAW
i I I i i
20 40 60 80 100 120
TIME, SEC
i
140
2.5
1.51
0.50
2O
VITHREGENERATION
40 60 80 100 120
TIME, SEC
t
140
6
"4
2
00
NO REGENERATION
60 80 I00 120 140
TIME, SEC
2.5
1.51
050
0
I ' , NO REGENERATION L_
20 40 60 80 100 120
TIME, SEC
140
#_L_ Rockwell InternationalSpace Systems Division
155
SSME TVC ACTUATOR INPUT POWER DURING ASCENT
STS-53SECOND STAGE ASCENT SIMULATION, NOMINAL MISSION
SSlVIE#1 PITCHSSME#I YAW
_ w,,.REENERAT'O"
0 IOO 200 300 400 500
TIME, SEC
600
0.35
s °°__o°_
00 I00
WITH REGENERATION
200 300 400 500
TIME, SEC
i
6O0
_. 5i 432
NO REGENERATION
I ,,_h i ... JILL ._--_,wJ1
0 loo 200 300 400 500
TIME, $EC
600
0.35
.o°2_0.2
0
NO REGENERATION
100 2OO 300 400 500
TIME, SEC
i
6O0
#i_ Rockwell InternationalSpace Systems Division
156
SSME TVC ACTUATOR INPUT POWER DURING ASCENT
STS-53SECOND STAGE ASCENT SIMULATION, NOMINAL MISSION
SSME#2 PITCHSSME#2 YAW
41
LW'TH EOENE 'ONj_1 | I =
0 100 200 300 400 500
TIME, SEC
3.5
o. 1.5
0.50
0
WITH REGENERATION
100 2O0 300 4O0 5OO
TIME, SEC
I
4
_2z.
00
I NO REGENERATION
100 200 300 400
TIME, SEC
500 6OO
o.i0 100
NO REGENERATION
200 300 400 500
TIME, SEC
l
600
#_L_ Rockwell InternationalSpace Systems Division
157
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SSME#1 PITCH SSME#1 YAW
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#_L_ Rockwell InternationalSpace Systems Division
159
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#dL_ Rockwell InternationalSpace Systems Division
162
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SSME TVC ACTUATOR INPUT ENERGY DURING ASCENT
STS-53SECOND STAGE ASCENT SIMULATION. NOMINAL MISSION
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TIME, SEC
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#_L_ Rockwell InternationalSpace Systems Division
163
SSME TVC ACTUATOR INPUT ENERGY DURING ASCENT
STS-53SECOND STAGE ASCENT SIMULATION. NOMINAL MISSION
SSME#3 PITCHSSME#3 YAW
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_L_ Rockwell InternationalSpace Systems Division
164
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POWER AND ENERGY REQUIREMENTS FOR THE MPS TVC ELA SYSTEMS
SSME #2 PITCH
STS-1 FLIGHT DATA
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II-" |
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# Lq Rockwell InternationalSpace Systems Division
166
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OI-20 SAILTESTDATA, NOMINAL ENTRY FROM ENTRYINTERFACETO END OF ROLLOUT
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10.08.06.04.02.0
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TIME, SEC
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TIME, SEC
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i i I
5O0 15001000
TIME, SEC
2O00
#_1_ Rockwell InternationalSpace Systems Division
168
9
POWER AND ENERGY REQUIREMENTS FOR THE ELEVON ELA SYSTEMS
OI-20 SAIL TESTDATA, NOMINAL ENTRY FROM ENTRYINTERFACETO END OF ROLLOUT
LEFTOUTBOARD ELEVON RIGHTOUTBOARD ELEVON
++=I "'+'1.2o _lj1.0 ..
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TIME, SEC
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0.00-0.01 ....
-- 0 500 1000 1500 2000
TIME, SEC
#I_ Rockwell InternationalSpace Systems Division
169
POWER AND ENERGY REQUIREMENTS FOR THE RUDDER AND SPEEDBRAKEELA SYSTEMS
O1-20SAILTESTDATA, NOMINAL ENTRY FROM ENTRY INTERFACE TO END OF ROLLOUT
RUDDER SPEEDBRAKE
,o I0.5
_O 0.0 fl _..-0.5
-I.0
' ' ooo-1.5 500 1000 1500 2
TIME, SEC
15.0 1. 10.OJ
,5.0I
0.0
I-I00 -5.0
"0
i t
500 1000
TIME, SEC
L ' '4
1500 2O00
2.0E-031.5E-(]3
1,0E-03
0.0E+005.0E-04
__ -5.0E-040
NO REGENER_,
WITH REGENERATION_
500 1000 1500 2000
TIME, SEC
0.07 I
0.06
0.050.040.030.02
'_ 0.010.00
Z -0.01
0
NO REGENERATION
WITH REGENERATION
500 1000 1500 2[]00
TIME,$EC
_dLq Rockwell InternationalSpace Systems Division
170
POWER AND ENERGY REQUIREMENTS FOR THE BODYFLAP ELA SYSTEM
OI-20 SAIL TESTDATA, NOMINAL ENTRY FROM ENTRYINTERFACETO END OF ROLLOUT
BODYFLAP
1.51
- 1.0
0.5
0.0
_-0.5
-I.0500 I000 1500 2000
TIME, SEC
3.0E-032.5E-03
2.0E-031.5E-03
_. 1.0E-03
5.0E-040.0E+00-5.0E-04
0
NO REGENERATION
WITH REGENERATION _S u-r--NL__
.J L___ _-
ii
5OO 1000 1500
TIME, SEC
2[]O0
_ Rockwell InternationalSpace Systems Division
171
_LI.
UOlSIAIO stuels,( S eolld$
ImJOflmJ_]UlIleM_.)IB _T_
S30EII'IOS 1:13MOd V'13 UO..-ISIN3N31::llnO=II::I AgEI3N30NV 143MOd 0"9
POWER AND ENERGY REQUIREMENTS FOR ELA POWER SOURCES
• THE SUBJECT POWER AND ENERGY REQUIREMENTS ARE NEEDED IN ORDER TODESIGN THE POWER SOURCES FOR THE ELA EFFECTOR SYSTEMS IN THE ORBITERAND EACH OF THE TWO SRB'S. A POWER SOURCE IS A SYSTEM WHICH MAYCGNTAIN REDUNDANT ELEMENTS SUCH AS BATTERIES, FUEL CELLS, ETC. TOPROVIDE ELECTRICAL POWER FOR THE ELA EFFECTOR SYSTEMS.
• THE REQUIREMENTS INCLUDE: (1) PEAK POWER AND THE DURATION OF THE PEAK,
(2) AVERAGE POWER AND THE OPERATING DURATION, (3) TOTAL ENERGY FOR THEMISSION, AND (4) POWER AND ENERGY DEMANDS AS FUNCTIONS OF MISSION TIME
(DUTY CYCLES).
• THE METHODS, DATA AND RESULTS FROM THE CALCULATION OF THESE REQUIRE-MENTS ARE PRESENTED IN THIS SECTION AS FOLLOWS.
• METHODS: SEE CHART 174.
• RESULTS: SEE CHARTS 175 THROUGH 180.
Rockwell InternationalSpace Systems Oivislon
173
METHOD FOR CALCULATION OF POWER SOURCE REQUIREMENTS
THE POWER AND ENERGY REQUIREMENTS FOR ELA POWER SOURCE IN THE ORBITER(OR IN EACH ONE OF THE TWO SRB'S) ARE CALCULATED AS FOLLOWS.
• POWER REQUIREMENTS
POWER REQUIREDFOR ELA POWERSOURCE IN ORBITER
(SRB)
m
SUM OF INPUT POWERREQUIREMENTS OF ELAEFFECTOR SYSTEMS IN THEORBITER (SRB) VERSUSMISSION TIME
• ENERGY REQUIREMENTS
ENERGY REQUIRED FORELA POWER SOURCE INORBITER (SRB)
mm
SUM OF INPUT POWERREQUIREMENTS OF ELAEFFECTOR SYSTEMS IN THEORBITER (SRB) VERSUSMISSION TIME
Rockwell Internal onalSpace Systems Division
174
CALCULATION RESULTS: POWER AND ENERGY REQUIREMENTS FORELA POWER SOURCES
ELA
POWER
SOURCE
ORBITER
EACH SRM
PEAK POWER, KW
(DURATION, SECONDS)
POWER REQUIREMENTS
AVERAGE POWER
(DURATION, MINUTES)
ASCENT ENTRY ASCENT ENTRY
1st STAGE 2nd STAGE 1st STAGE ._ndSTAGE18 (1) 19 (1) 19 (1) 0.05 (2) 0.01 (6.4) 0.02 (31)
13 (1)0.01 (2)
-RGY
_ED
THE
0.74
0.02
NOTES: (1) THE POWER AND ENERGY REQUIREMENTS FOR THE ELA EFFECTOR SYSTEMSDO NOT INCLUDE THOSE FOR THE EXISTING SHU'I-I'LE AVIONICS. FOR EXAMPLE,THE EXISTING ORBITER AVIONICS REQUIRE AN AVERAGE POWER OF ABOUT
11.3 KILOWATrS (KW).
(2) THE ABOVE REQUIREMENTS ARE INDEPENDENT OF THE REDUNDANCY OF THEPOWER SOURCE. FOR EXAMPLE, EACH BATI'ERY OF A QUAD-REDUNDANT (FAILOPERATIONAL/FAIL SAFE) POWER SOURCE WITH FOUR BATI'ERIES SHALL MEET50 PERCENT OF THE REQUIREMENTS. (100 PERCENT FOR REMAINING TWO
BATTERIES)
#_L_ Rockwell InternationalSpace Syslems Division
175
TOTAL POWER AND ENERGY REQUIREMENTS FOR THE SSME-TVC ELA SYSTEMS
FIRST STAGE, 3 SIGMA DISPERSION
_. 15
=_ 10
-- 00 20 40 60 80 100 120
TIME, SEC
"4
140
0. ! T NO REGENERATION
0.( 3
o.,sI " _ "\0.04 | _ WITH REGENERATION
0 20 40 60 80 100 120 140
TIME, SEC
_4Lq Rockwell InternationalSpace Systems Division
176
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TOTAL POWER AND ENERGY REQUIREMENTS FOR THE MPS TVC ELA SYSTEMS
STS-1 FLIGHT DATA
30
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-15 ...........................................................................-400 -300 -200 -100 0 100 200 300 400 500
TIME. SEC
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Rockwell InternationalSpace Systems Division
178
TOTAL POWER AND ENERGY REQUIREMENTS FOR THE AEROSURFACE ELA SYSTEMS
20. 15l0
L 0
-5z -10
0
i J i
500 10130 1500 2000
TIME, SEC
i 0.250.2
i_ 0.15,,, 0.1
..z.o._-0.05
0
NO REGENERATION
WITH R____
500 1000 1500 2000
TIME, SEC
OI-20 SAIL Test Data, Nominal Entry
From Entry Interlace to End of Rollout
#_ Rockwell InternationalSpace Systems Division 179
TOTAL POWER AND ENERGY REQUIREMENTS FOR THE SRM ELA SYSTEMS
LEFT SRBSTS-56 Flight Simulation, 3 Sigma Dispersion
RIGHT SRB
10
-10 ' "0 20 40 60 80 140
TIME, SEC
,o
_o_I I I I I I
-6 20 40 60 80 100 120
TIME, SEC
0.02
0.015
joo,0.005i o
NO REGENERATION /
.0 50 100 150
TIME, SEC
r_ 0.02
,m,,-r- 0.015
zm_ 0.01'_ 0.005
o0
NO REGENERATIO__
WITH REGENERATIONi 1
50 100
TIME, SEC
I
140
I
150
#_L_ Rockwell InternationalSpace Systems Division
180
7.0 CONCLUSIONS AND RECOMMENDATIONS
_I_i,_ _ell Intemalionalscemsy_ms _
181
CONCLUSIONS
THE RESULTS OF THE STUDY INDICATE MAGNITUDES OF PEAK POWER, AVERAGE
POWER AND ENERGY WELL WITHIN THE PREVIOUS ESTIMATES AND EFFECTOR
DESIGN VALUES. THE SIMULATION AND FLIGHT DATA USED FOR THE ANALYSIS
WERE OBTAINED FOR SPECIFIC FLIGHTS. THE RESULTS SHOULD BE CONSIDERED
TO BE GENERALLY TYPICAL AND DO NOT NECESSARILY INCLUDE WORST CASE
MANEUVERS.
THE DATA SHOW VERY FEW BURSTS OF EFFECTOR INPUT POWER EXCEEDING
ABOUT 5 KW OF INPUT POWER. THE PEAK ELEVON POWER OUTPUT OCCURSDURING A SHORT PERIOD ON LANDING WITH FEW PEAKS EXCEEDING 2 KW DURING
FLIGHT PRIOR TO LANDING.
THE SPACE SHUTTLE EFFECTOR SYSTEMS TYPICALLY EXPERIENCE AIDING LOADS
AS WELL AS OPPOSING LOADS. THE ELA SYSTEM MUST THEREFORE BE CAPABLE
OF HANDLING THE ENERGY RETURNED TO THE SYSTEM BY THE AIDING LOADS AS
WELL AS PROVIDING THE REQUIRED POWER FOR THE OPPOSING LOADS.
Rockwell InternationalSpace Systems Division
182
RECOMMENDATIONS
PERFORM TESTS ON REPRESENTATIVE ELA SYSTEMS WITH AIDING AND OPPOSING
LOADS, FIXED LOADS AND FLIGHT LOAD PROFILES TO (1) DETERMINE THE EFFECTSOF THE ENERGY RETURNED TO THE ELA SYSTEM BY THE AIDING LOADS AND OF
THE FLIGHT PROFILE TIME CHARACTERISTICS ON THE ELA SYSTEM AND (2) OBTAIN
POWER AND ENERGY DATA TO CONFIRM OR UPDATE THE SYSTEM EFFICIENCIES
AND LOSSES.
PERFORM ADDITIONAL FLIGHT SIMULATION ANALYSES WITH FAILURES AND OFF-NOMINAL CONDITIONS TO EVALUATE WORST CASE POWER AND ENERGY DEMANDS
AND TO EVALUATE ANY EFFECT ON THE ELA LOSSES DUE TO THE FLIGHT PROFILE
CHARACTERISTICS.
Rockwell InternationalSpace Systems Division
183
4. _Ue am Subdtle
Strategic AvionicsSubtask 3 - IA, Electrical
System FinaiReport
7. Au_ods)
Ben T. F Lum
charles •Pond
William McDermott
Report Documentation Page
2. Government AcceS._on No.
Technology. Definition StudiesActuation.(ELA)
9. PeHo_ing Organization Name and Adore_
Rockwell International Corp.
Space Systems Division12214 Lakewood Blvd.
I2. S_o_g Agency Name and Address
National Aeronautics and
Johnson Space CenterHouston, Texas 77058
Space Administration
3. RedpienC's Catalog No.
5. Report DaP
September 30, 1993
6. Performing Organ_.ation Code
I 8. Performing Organ.Ezad°n Repo_ No.
IssD93Doso 10. Work Unit No.
11. Con_a_ or Grant No.
NAS9-18880
13. Type _f Repo_ and Peri_ Co_red
Final 1993
14. Sponsoring Agency Code
15. Supplementary Notes
16. Abstra_
Electrical Actuator (ELA) power efficiency and requirements are examined for
space system application. Requirements for Space Shuttle effector systems
are presented, along with preliminary ELA trades and selection to form a
preliminary ELA system baseline. Power and energy requirements for thisbaseline ELA system are applicable to the Space Shuttle and similar space
vehicles.
17. K_ Words (Suggested by Author(s)}
Electrical Actuation, ELA,
Mechanical Actuators, EMA,
Space Shuttle, effector
-19. Security Closer. (of _is report|
Unclassified
Electro-
Actuation,
18. Dis_ution Statement
Unclassified _ Unlimited
_. SecunW (_fo (Of th_ pa_} •
Unclassified
21. No, of pages
183
t 22. Price
JASA FORM 1626 OCT 86
, I