TEST REQUIREMENTS FOR SPACE VEHICLES - MIL-STD-188everyspec.com/MIL-STD/MIL-STD-1500-1599/download.php?spec=MI… · Test Requirements for Space Vehicles MIL-STD-1540B (USAF) 1. This

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MILITARY STANDARD

TEST REQUIREMENTS FOR SPACE VEHICLES

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MIL-STD-1540B (USAF)10 OC~BER 1982

Department of the Air Force

Washington, D. C. 20301

Test Requirements for Space Vehicles

MIL-STD-1540B (USAF)

1. This Military Standard is approved for use by theDepartment of the Air Force and is available for use by allDepartments and Agencies of the Department of Defense.

2. Beneficial comments (recommendations, additions, ordeletions) and any pertinent data which may be of use inimproving this document should be addressed to:

SD/AJd4Pd. BOX 92960Worldway Postal CenterLos Angeles, California 90009

by using the self-addressed Standardization DocumentImprovement Proposal (DD Form 1426) appearing at the end ofthis document or by letter.

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MIL-STD-1540B (USAF)10 OCTOBER 1982

E’ORWORD

All space vehicles are subjected to extensive ground testing toensure their successful operational use. This standardestablishes a uniform set of definitions and ground testingrequirements for space vehicles, whether launched by anexpendable launch vehicle or the recoverable Space Shuttle.Although this standard is not applicable to launch vehicles, itis recommended that the component test requirements be used forlaunch vehicle components to obtain the required highreliability. The test requirements specified are a composite ofthose tests currently used in achieving successful space mis-sions. It is intended that these test requirements should betailored to the specific space progrsm or project consideringdesign complexity, state of the art, mission criticality andacceptable risk.

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This is Document 0002b, Archive 0001b

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1.

1.11.21.3

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2.12.2

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3.13.23.33.43.53.63.73.83.93.103.113.123.133.143.153.163.173.183.193.203.213.223.233.243.25

3.263.273.283.293.303.31

TABLE OF CQNTENTS

PageSCOPE ................................................

Purpose .............................................Application ........................................Classifications ....................................,

REFERENCED 00CUMEt4T5.................................

Issues of Documents .................................Order of Precedence .................................

DEFINITIONS ..........................................

Acceptance Tests ....................................Airborne SuppOrtquiPment (ACE)....................Ambient Environment .................................Burst Pressure ......................................Component ...........................................Computer Program component (OK) ....................Decibel (dB)........................................Design Environments, Space Vehicle ..................Design Environments, Space Vehicle Components .......Development Tests ...................................Development Test Vehicle (TV)................... ...Environmental Design Margin .........................Flight Unit .........................................Item Levels..........................................Launch System .......................................Launch Vehicle, Expendable ..........................Iaunch Vehicle, Recoverable............. ............Limit Iaad.....:.,...................................Maximum Predicted Acceleration ......................Maximum Predicted Acoustic Environment ..............Maximum Predicted Operating Pressure ................Maximum Predicted Pyro Shock ~vironment ............maximum Predicted Random Vibration Environment ......Maximum Predicted Sinusoidal Vibration Environment..Maximum and Minimum Predicted Component

Temperatures ......................................Moving Mechanical Assemblies. .......................Multipacting ........................................On-Orbit System .....................................Operational Modes ...................................Optional Tests ......................................Part ................................................

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:66777999999101010101111

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MIL-STD-1540B (USAF)10 OC’IX3BER1982

3.323.333.343.353.363.373.383.393.403.413.423.433.443.453.46

4.

4.14.1.14.1.24.1.34.1.44.1.54.24.2.14.2.24.2.34.34.4

5.

5.15.25.3

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6.16.1.16.1.26.26.2.16.2.2

6.2.36.2.46.2.5

Pascal ..............................................Prelaunch Validation Tests ..........................Proof Pressure ......................................Qualification Tests .................................Qualification TestVehicle. .........................Service Life ........................................Space Vehicle .......................................Subassembly .........................................Subsystem ...........................................System ..............................................Test Discrepancy ....................................Test Environments ...................................Test Unit Failures ..................................Thermal Uncertainty Margin ..........................Ultimate Laad .......................................

GENERAL REQUIREMENTS .................................

Testing Philosophy ..................................Development Testing ................................Qualification Testing ..............................Acceptance Testing .................................Prelaunch Validation Testing. ......................Other Testing ......................................

Test Plans and Procedures. ..........................Preparation ........................................Test Condition Tolerances ..........................Tailoring ..........................................

Retest ..............................................Test Data Analysis ..................................

DEVELOPMENT TESTING .................................,

General ..............................................Component Development Tests ................ .........Space Vehicle and Subsystem Development Testing .....

QUALIFICATION TESTING... .............................

General Qualification Test Requirements .............Qualification Hardware .............................Qualification Test keels ..........................

Space Vehicle Qualification Tests...................Functional Test, Space Vehicle Qualification .......Electromagnetic Compatibility Test, Space VehicleQualification ....................................

Acoustic Test, Space Vehicle Qualification .........Vibration Test, Space Vehicle Qualification ........Pyro Shock Test, Space Vehicle Qualification .......

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6.2.66.2.76.2.86.2.96.36.3.1

6.46.4.16.4.26.4.36.4.46.4.56.4.66.4.76.4.86.4.96.4.106.4.116.4.126.4.136.5

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7.17.1.17.1.27.1.37.1.47.1.57.1.67.1.77.1.87.1.97.27.37.3.17.3.27.3.37.3.47.3.57.3.67.3.77.3.87.3.97.4

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Pressure Test, Space Vehicle Qualification .........Thermal Vacuum Test, Space Vehicle Qualification. ..Thermal Balance Test, Space Vehicle Qualification. .Thermal Cycling Test, Space Vehicle Qualification. .Subsystem Qualification Tests.......................Structural Static Load Teat, Subsystem

Qualification ....................................Component Qualification ‘rests.......................Functional Test, Component Qualification ...........Thermal Vacuum Test, Component Qualification .......Thermal Cycling Test, Component Qualification ......Sinusoidal Vibration Test, Component Qualification.Handom Vibration Test, Component Qualification .....Acoustic Test, Component Qualification .............Pyro Shock Test, Component Qualification ...........Acceleration ~st, Component Qualification .........Humidity Test, Component Qualification .............Pressure Test, Component Qualification .............Leakage Test, Component Qualification ..............EMC Test, Component Qualification ..................Life Test, Component Qualification .................

Subassembly Level Qualification Tests ...............

ACCEPTANCE TESTS .....................................

Space Vehicle Acceptance Tests ......................Functional Test, Space Vehicle Acceptance ..........R4C Test, Space Vehicle Acceptance .................Acoustic Test, Space Vehicle Acceptance ............Vibration Test, Space Vehicle Acceptance ...........Pyro Shock Test, Space Vehicle Acceptance ..........Pressure Test, Space Vehicle Acceptance .............Thermal Vacuum Test, Space Vehicle Acceptance ......Thermal Cycling Teat, Space Vehicle Acceptance.. ...Storage Tests, Space Vehicle Acceptance ............

Subsystem Acceptance Tests ..........................Component Acceptance Tests ..........................Functional Performance Test, Component Acceptance..Thermal Vacuum Test, Component Acceptance ..........Thermal Cycling Test, Component ACCeptaIICe . . . . . . . . .

Random Vibration Test, Component Acceptance ........Acoustic Test, Component Acceptance ................Pyro Shock Screening Test, Component Acceptance ....Pressure Test, Component Acceptance ................Leakage Test, Component Acceptance .................Burn-In Test, Component Acceptance .................

Subassembly Level Acceptance Tests ..................

FLIGHT USE OF QUALIFICATION EQUIPMENT. ...............

3536384041

41434445474849505152535456585960

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61626263636464646566666668686869707071727274

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8.18.1.18.1.28.2

8.2.18.2.28.38.4

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9.19.29.39.49.4.19.4.29.4.39.4.4

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I.-11.

IIII.IV.

Use of theComponentComponent

Use of the

Qualification Components for Flight ...... 75 ●Qualification Tests ...................... 76Certification for Flight ................. 76Fliqht Vehicle for SDace Vehicle Level

Qualification ..................................... 77Space Vehicle Qualification Tests .................. 77Space Vehicle Certification for Flight ............. 77Use of the Qualification Vehicle for Flight ......... 78Other ............................................... 78

PRELAUNCH VALIDATION TESTS ........................... 79

General Requirements ................................ 79Prelaunch Validation Test Flow ...................... 79Prelaunch Validation Test Configuration ............. 80Prelaunch Validation Test Descriptions .............. 80Functional Test .................................... 80Propulsion System Leak and Functional Test ......... 81Integrated System Tests ............................ 81Compatibility Test, On-Orbit System ................ 81

NOTES ................................................ 83

TABLES

Space Vehicle Qualification Tests.................... 30Component Qualification Tests........................ 44Space Vehicle Acceptance Tests ....................... 61Component Acceptance Tests ........................... 67

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SECTION 1

SCOPE

1.1 PURPOSE

~’ This standard establishes uniform definitions,environmental criteria, test requirements, and test methods for

(. space vehicles and their subsystems and components.

1.2 APPLICATION

The tailored application of these test requirements to aparticular space program is intended to assure a high level ofconfidence in achieving a successful space mission. This

I standard is intended for use in the procurement of apace vehiclehardware, including space vehicles and airborne supportequipment that remain in the Space Shuttle Orbiter duringorbital flight, aa well as orbital satellites.I

1.3 CLASSIFICATIONS

o The testsqualification,

specified herein are classified as development,acceptance, prelaunch validation, or other tests.

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MIL-STD-1540B (USAF)10 OCl_OBER 1982

SECTION 2

REFERENCED DOCUMENTS

2.1 ISSUES OF DOCUMENTS

The following documents of the issue in effect on the dateof invitation for bids or request for proposal form a part ofthis standard to the extent specified herein.

Military Standards

MIL-STD-1541 (USAF) Electromagnetic CompatibilityRequirements for Space Systems

(Copies of specifications, standards, and publications requiredby contractors in connection with specific procurement functionsshould be obtained from the contracting officer or as directedby the contracting officer. )

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2.2 OROER OF PRECEDENCE

In the event of conflict between documents referencedherein and the contents of this standard, the contents ofstandard shall be considered the superseding requirement.

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SECTION 3

DEFINITIONS

3.1 ACCEPTANCE TESTS

Acceptance tests are the required formal tests conducted todemonstrate acceptability of an item for delivery. They areintended to demonstrate performance to specification .requirements and to act as quality control screens to detectdeficiencies of workmanship, material, and quality.

3.2 AIRBORNE SUPPORT EQUIPMENT (ASE~

Airborne support equipment is the equipment installed in arecoverable launch vehicle to provide support functions andinterfaces for the space vehicle during launch and orbitaloperations of the recoverable launch vehicle. This includes thehardware and software which provides the structural, electrical,electronic, and mechanical interfaces with the launch vehicle.ASE is recovered with the launch vehicle.

3.3 AMBIENT ENVIRONMENT

Ambient environment is defined as normal room conditionswith temperature at 23 ~ 10 deg C (73 deg ~ 18 deg F),atmospheric pressure 101 + 2.0/-23 kilopascals (29.9 + 0.6/-6.8in. Xg) and relative humidity of 50 ~ 30 percent.

3.4 BURST PRESSURE

The burst pressure is the maximum test pressure thatpressurized components withstand without rupture to demonstratethe adequacy of the design in a qualification test. It iS equalto the product of the maximum expected operating pressure, burstpressure design factor, and a factor corresponding to thedifferences in material properties between test and designtemperatures.

3.5 apDNENT

A component is a functional unit that is viewed as anentity for purposes of analysis, manufacturing, maintenance, orrecord keeping. Examples are hydraulic actuators, valves,batteries, electrical harnesses, and individual electronic boxessuch as transmitters, receivers, or multiplexer.

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3.6 COMPUTER PROGRAM COMPONENT (CPC)

A computer program component (CPC) is a functionally orlogically distinct part of a computer program configuration item(CPCI) that is distinguished for purpose of convenience indesigning and specifying acomplex CPCI as an assembly ofsubordinate elements.

3.7 DECIBEL (dB)

Amplitude and power ratios may be expressed in decibels(dB). Ten times the logarithm to the base 10 of the ratio ofone power level to a reference power level is the value of thepower, or power ratio, in dB. For pressures, electricalcurrents, voltages, sinusoidal vibration, and other parameters,the amplitude rather than power is frequently measured. Twentytimes the logarithm to the base 10 of the ratio of one amplitudemeasurement to a reference amplitude is the value of the powerratio in dkl. The referenced amplitude or referenced power levelmust be clear. For sound pressure levels the amplitude of thereference pressure is 0.0002 dynes per square centimeter.

3.8 DESIGN ENVIRONMENTS, SPACE VEHICLE

The design environments for a space vehicle are the ocomposite of the various environmental stresses to which theSpa;e vehicle must be designed. Each of the design environmentsfor a space vehicle is based upon:

a. The maximum and minimum predicted environmentsduring the operational life of thespace vehicle,plus

b. An environmental design margin. (see 3.12) thatincreases the environmental range to provide anacceptable level of confidence that a failurewill not occur during the service life of thespace vehicle.

3.9 DSSIGN ENVIRONMENTS, SPACE VEHICLE COMPONENTS

The design environments for space vehicle components arethe composite of the various environmental stresses to which thespace vehicle hardware components must be designed. Each of thedesign environments for a space vehicle component is based upon:

a. The maximum and minimum predicted environmentsduring the operational life of the component, orfor temperature, a standard thermal range between

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MIL-sTD-1540B (USAF)10 OCTOBER 1982

0-24 deg C and +61 deg C when the predicted rangeis less severe, plus

b. h environmental design margin (see 3.12) thatincreases the environmental range to provide anacceptable level of confidence that a failurewill not occur during the service life of thecomponent (see 3.37).

3.10 DEVELOPMENT TESTS

Development tests include all tests conducted to obtaininformation to aid in the design and manufacturing processes.Development tests are conducted to generate design parameters,validate design concepts, verify design criteria, determinedesign margins, identify failure modes, and to verifymanufacturing processes. Development testing may be informal inthat controlled design and test documentation, formalcertification, formal retest requirements, and flight typehardware are usually not required.

3.11 DEVE~PMENT TEST VEHICLE (DTV)

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A development test vehicle (DTV) iS a SpaCe vehicle orsubsystem fabricated to provide engineering design and testinformation concerning the validity of analytic techniques andassumed design parameters, to uncover unexpected system resPonsecharacteristics, to evaluate design changes, to determineinterface compatibility, to prove test procedures andtechniques, and to determine if the equifxsentmeets itsperformance specifications. DTVS include engineering testmodels, thermal models, and structural models.

3.12 ENVIRONMENTAL DESIGN MARGIN

An environmental design margin for an item is an increasein the environmental range used for the design (and for thequalification testing) of an item to reduce the risk of anoperational failure. It may include increases in the maximumlevels, decreases in the minimum levels, and increases in thetime exposure to the extreme levels. The environmental designmargin ia intended:

a. TO accommodate differences among qualificationand flight units due to variations in parts,materiala, processes, manufacturing, testing, anddegradation during useage;


MIL-STD-1540B (USAF)10 OClilBER1982

b. To incorporate the allowable test conditionstolerances; “

c. ‘lbavo,idqualification test levels that are lesssevere than the acceptance test ranges oroperating ranges;

d. To help assure against fatigue failures due torepeated testing and operational use.

Unless otherwise specified, the test condition tolerancesallowed by this standard are assumed to be incorporated in theenvironmental design margin. For example, space vehicle itemsare designed, unless otherwise specified, to thermalenvironments 10 deg C higher and 10 deg C lower than the maximumpredicted thermal ranges, (see 3.25). This 10 deg Cenvironmental design margin includes a ~ 3 deg C tolerance foracceptance test conditions and a ~ 3 deg C tolerance forqualification test conditions.

Unless otherwise specified, space vehicle items are alsodesigned to acoustic noise and random vibration environmentsthat are 6 dB above the maximum predicted levels. This 6 dBenvironmental design margin for acoustic noise and randomvibration includes a ~ 1.5 dB tolerance in the overall level(integrated root mean square value over the total frequencyrange Of the test spectrum) for acceptance test conditions and~ 1.5 dB tolerance for qualification test conditions.

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When the qualification or acceptance tests are controlledusing test condition tolerances with magnitudes less thanspecified herein (3 deg C or 1.5 dB), the environmental designmargins (10 deg C or 6 dB) may be reduced accordingly. Forexample, if qualification and acceptance acoustic tests wereboth controlled to ~ 1.0 dB, the design margin would be 5 dBinstead of 6 dB. If larger test condition tolerances areallowed, then the design margins would be increased accordingly.

Other environmental design margins applicable to spacevehicle items include 6 dB for shock, 6 dB for sinusoidalvibration, a factor of 2 for launch or injection acceleration,and a factor of 1.25 for maximum acceleration of deployedcomponents on a spinning space vehicle.

Another element of the environmental design margin is thetime the item is exposed to the design environmental levels. Mincrease in exposure time or number of cycles over that expectedin operation is usually specified for vibration and acousticdesign environments to increase confidence that wearout or


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fatigue failures will not occur. Of course, the environmentaldesign margins may be changed to either higher or lower levels,or to longer os shorter exposure times, depending upon specificprogram requirements and allowable risk.

3.13 FLIGHT UNIT

A flight unit is a component, subsystem, orintended for actual flight.

3.14 IT~ LEVELS

The item levels used in this standard, from

space vehicle

the simplestdivision to the more complex, are: part, subassembly,component, subsystem, space vehicle, and system.

3.15 LAUNCH SYSTEM

A launch syetem is the composite of equipment, skills, andtechniques capable of Launching and boosting the space vehicleinto orbit. The launch system includes the space vehicle, thelaunch vehicle, and related facilities, equipment! material,software, procedures, services, and personnel required for theiroperation.

3.16 LAUNCH VEHICLE, EXPENDABLE

An expendable launch vehicle ia a composite of the boosterinitial stages, injection stages, space vehicle adapter, andfairing having the capability of a single launching andinjection of a space vehicle or vehicles into orbit.

3.17 LAUNCH VEHICLE, RECOVERABLE

A recoverable launch vehicle is a composite of boosterstages, injection stages, and a space vehicle carrier having thecapability to carry a space vehicle to a parking orbit.ElementS of the launch vehicle may be recovered for use and mayreturn to earth with or without the space vehicle.

3.18 LIMIT ~AD

The limit load is the maximum anticipated load, orcombination of loads, which a structure may be expected toexperience during the performance of specified missions inspecified environments. Since theexperienced in service are in partmethods for predicting limit loadsappropriate.

actual loads that arerandom in nature, statisticalare employed wherever



3.19 MAXIMUM PREDICTED ACCELERATION

The maximum predicted acceleration is the accelerationvalue determined from the combined effects of the quasi-steadyacceleration and the transient response Of the vehicle to engineignition, engine burnout, and stage separation. Where thenatural frequency of the component mount or mounting Structuremay couple with these engine-initiated transients, the maximumpredicted acceleration level shall account for the possibledynamic amplification. Different levels may be applied indifferent axes if the mounting orientation in the space vehicleis specified.

3.20 MAXIMUM PREDICTED ACOUSTIC ENVIRONMENT

The maximum predicted acoustic environment is the extremevalue of fluctuating pressure occurring on the external surfaceof the space vehicle which occurs during liftoff, poweredflight, or reentry. The maximum predicted acoustic environmenttest spectrum is specified based on one-third octave bands overa frequency range of 32 to 10,000 hertz (Hz). The duration ofthe maximum environment is the total period when the overallamplitude is within 6 dB of the maximum overall amplitude.Where sufficient data are available, the maximum predictedenvironment may be derived using parametric statisticalmethods. The data must be tested to show a satisfactory fit tothe assumed underlying distribution. The maximum predictedenvironment is defined as equal to or greater than the value atthe ninety-fifth percentile value at least 50 percent Of thetime. Where there are less than three data samples, a minimummargin of 3 dB is applied to the prediction to account for thevariability of the environment.

3.21 MAXIMUM PREDICTED OPERATING PRESSURE

The maximum predicted operating pressure is the workingpressure applied to a component by the pressurizing system withthe pressure regulators and relief valves at their upperoperating limit, including the effects of temperature, transientpeaks, and vehicle acceleration.

3.22 MAXIMUM PREDICTED PYRO SHOCK ENvIRONMENT

The pyro shock environment imposed on the space vehiclecomponents is due to structural response when the space orlaunch vehicle electro-explosive devices are activated.Resultant structural response accelerations resemble the form ofsuperimposed complex decaying sinusoids which decay to a fewpercent of their maximum acceleration in 5 to 15 milliseconds.

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The maximum predicted pyro shock environment is specified as amaximum absolute shock response spectrum determined by theresponse of a nu~er of single-degree-of-freedom systems usingQ = 10. The Q is the acceleration amplification factor at theresonant frequency for a lightly damped system. This shockresponse spectrum is determined at frequency intervals ofone-sixth octave or less over a frequency range of 100 to 10,000Hz. Where sufficient data are available, the maximum predictedenvironment may be derived using parametric statisticalmethods. The data must be tested to show a satisfactory fit tothe assumed underlying distribution. The maximum predictedenvironment is defined as equal to or greater than the valbe atthe ninety-fifth percentile value at least 50 percent Of thetime. Where there are less than three data samples, a minimummargin of 4.5 dB is applied to account for the variability ofthe environment.

3.23 MAXIMUM PREDICTED RANOOM VIBRATION ENVIRONMENT

The random vibration environment imposed on the spacevehicle components is due to the lift-off acoustic field,aerodynamic excitations, and transmitted structure-bornevibration. The maximum predicted random vibration environmentis specified as a power spectral density, based on a frequencyresolution of 1/6 octave (or narrower) bandwidth analysis, overa frequency range of 20 to 2000 Hz. A different spectrum may berequired for different equipment zones or for different axes.The component vibration levels are based on vibration responsemeasurements made at the component attachment points duringground acoustic tests or during flight. The duration of themaximum environment ia the total period during flight when theoverall amplitude is within 6 dB of the maximum overallamplitude. Where sufficient data are available, the maximumpredicted environment may be derived using parametricstatistical methods. The data must be tested to show asatisfactory fit to the assumed underlying distribution. Themaximum pre~icted environment is defin;d ~s equal to or greaterthan the value at the ninety-fifth percentile value at least 50percent of the time. Where there are less than three datasamples, a minimum margin of 3 dB is applied to account for thevariability of the environment.

3.24 MAXIMUM PREDICTED SINUSOIDAL VIBRATION ENVIRONMENT

The sinusoidal vibration environment imposed on the spacevehicle subsystems and components is due to sinusoidal andnarrow band random forcing functions within the launch vehicleor space vehicle during flight or from ground transportation andhandling. In flight sinusoidal excitations may be caused by

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MIL-STD-1540B (uSAP)10 OCTOBER 1982

unstable combust.frequencies with

on, by coupling of structural resonantpropellant system resonant frequencies (pOGO)r

or by imbalances in rotating equipnent in the launch vehicle orspace vehicle. Sinusoidal excitations may occur during groundtransportation and handling due to the resonant response oftires and suspension systems of the transporter. The maximumpredicted sinusoidal vibration environment is specified over afreqUWICy range of 20 to 2000 Hz for flight excitation and 0.3to 300 Hz for ground transportation excitation. Wheresufficient data are available, the maximum predicted environmentmay be,derived using parametric statistical methods. The datamust be tested to show a satisfactory fit to the assumedunderlying distribution. The maximum predicted environment isdefined as equal to or greater than the value at the. .ninety-fltth percentile value at least 50 percent of the time.Where there are less than three data samples, a minimum margin.of 3 dB is applied to account for the variability of theenvironment.

3.25 MAXIMUM AND MINIMUM PREDICTED COMPONENT TEMPERATURES

The maximum and minimum predicted component temperaturesare the highest and lowest temperatures that can be expected tooccur on each component of the space vehicle during all

I operational modes plus an uncertainty factor. The componenttemperatures are predicted by an analytical thermal model forall operational modes. This analytical model includes theeffects of worst case combinations of equipment operation,internal heating, space vehicle orientation, solar radiation,eclipse conditions, ascent heating, and degradation of thermal

I surfaces during the life of the mission. The analytical modelused in this prediction is usually validated by a space vehicle

I thermal balance test under the worst case operational modes. ~appropriate mar9in for uncertainties is’applied to the extremecomponent temperatures predicted by the analytical model, even

I after validation by a thermal balance test, to obtain themaximum and minimum predicted temperatures. This marginaccounts for uncertainties in parameters such as complicatedview factors, surface properties, contamination, radiationenvironment, joint conduction, and inadequate ground

I simulation. Because of these uncertainties, an uncertaintymargin (see 3.45) of at least 11 deg C is included in all casesin determining the maximum or minimum predicted temperatures forspace vehicle components. This 11 deg C thermal margin isapplied to the temperature predictions made after thequalification thermal balance test. This implies that evenlarger thermal margins are required at the beginning of aprogram to accommodate changes that typically evolve from

I preliminary design to the final product.

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3.26 MOVING MECHANICAL ASSEMBLIES

Moving mechanical assemblies are the mechanical orelectromechanical devices that control the movement of onemechanical part of a space vehicle relative to another part.They include but are not limited to deployment mechanisms,pointing mechanisms, drive”mechanisms, despin mechanisms, andthe actuators, motors, linkages, latches, clutches, springs,cams, dampers, booms, gimbals, gears, bearings, andinstrumentation that are an integral part of these mechanicalassemkdies.

3.27 MULTIPACTING

Multipacting is the resonant back and forth flow ofsecondary electrons in a vacuum between two surfaces separatedby a distance such that the electron transit time is an oddintegral multiple of one half the period of the alternatingvoltage impressed on the surfaces. Multipacting requires anelectron impacting one surface to initiate the action, andrequires the secondary emission of one or more electrons at eachsurface to sustain the action. Multipacting is an unstableself-extinguishing action which can occur at pressures less than6.65 pascals (0.05 ‘rorr),excePt that it maY become stable atpressures less than 0.0133 pascals (0.0001 ‘Ibrr). The pittingaction resulting from the secondsry emission of electronsdegrades the impacted surfaces. The secondary electron emissioncan also increase the pressure in the vicinity of the surfacescausing ionization (corona) breakdown to occur. These effectscan cause degradation of performance or permanent failure of theradio frequency cavities, waveguides, or other devices involved.

3.28 ON-ORBIT SYSTEM

An on-orbit system is the composite Of equipnent, skills,and techniques permitting on-orbit operation of the spacevehicle. The on-orbit system includes the space vehicle(s), thecommand and control network, snd related facilities, equipment,material, software, procedures, services, and personnel requiredfor their operation.

3.29 OPERATIONAL MODES

The operational modes for a component or space vehicleinclude all combinations of operational configurations orconditions that can occur during their service life (see 3.37).Some examples are: power on or power off, command modes,readout modes, attitude control modes, antennas stowed or

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deployed, and spinning or despun.

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3.30 OPTIONAL TESTS

Optional tests are those tests which are not normallyrequired, but which may be necessary due to special requirementsof usage or to peculiarities of a particular configuration.

3.31 PART

A part is a single piece, or two or more pieces joinedtogether, which are not normally subject to disassembly withoutdestruction or impairment of the design use. Some examples areresistors, transistors, integrated circuits, relays, capacitors,gears, screws, and mounting brackets.

3.32 PASCAL

The pascal (Pa) is the unit of pressure in theinternational metric system. One pascal is equal to 0.000145lbs per sq in. or 0.0075 Torr.

~3.33 PRELAUNCH VALIDATION TESTS

Prelaunch validation tests are conducted at the launch baseto assure readiness of the hardware, computer programs,personnel, and procedures to support launch and the program omission.

I3.34 PROOF PRESSURE

The proof pressure is the test pressure that pressurizedcomponents can sustain without detrimental deformation. Theproof pressure is used to give evidence of satisfactoryworkmanship and material quality, or to establish maximumpossible flaw sizes. It is equal to the product of maximum

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expected operating pressure (see 3.21), proof pressure designfactor, and a factor accounting for the difference in materialproperties between test and design temperature.

I 3.35 QUALIFICATION TESTS

Qualification tests are formal contractual demonstrationsthat the design, manufacturitig, and assembly have resulted inhardware and computer programs conforming to specificationrequirements.

3.36

used

QUALIFICATION TEST VEHICLE

A qualification test vehicle is a space vehicle which isfor qualification tests. The qualification test vehicle

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MIL-ST13-1540B (USAF’)10 OCTOBER 1982

should be representative of the flight hardwarenecessary to-achieve a valid qualification testbeing qualified.

3.37 SERVICE LIFE

to the extentfor the items

The service life of a component or space vehicle is thetotal life expectancy of the item. The service life starts atthe completion of assembly of the item and continues through allacceptance testing, handling, storage, transportation, launchoperations, orbital operations, refurbishment, retesting,reentry or recovery from orbit, and reuse that may be requiredor specified for the item.

3.38 SPACE VSHICLE

A space vehicle is a complete, integrated set of subsystemsand components capable of supporting an operational role inspace. A space vehicle may be an orbiting vehicle, a majorportion of an orbiting vehicle, or a payload which performs itsmission while attached to a recoverable launch vehicle. Theairborne support equipment which is peculiar to programsutilizing a recoverable launch vehicle shall be considered a

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part of the space vehicle being carried by the launch vehicle.

3.39 S~A5S5MBLY

The term subassembly denotes two or more parts joined

1-

together to form a stockable unit which is capable ofdisassembly or part replacement. Sxamples are a printed circuitboard with parts mounted, or a gear train.

3.40 SUBSYSTE?!

A subsystem is an assembly of two or more components,including the supporting structure to which they are mounted,and any interconnecting cables or tubing. A subsystem iscomposed of functionally related components that perform one ormore prescribed functions. TvPical sDace vehicle subsystems are.-electrical power, attitude control, t&emetry, instrumentation,command, structure, thermal control, and propulsion.

3.41 SYSTEM

A system is the composite of equipnent, skills, andtechniques capable of performing or supporting an operational “role. A system includes all operational equipnent, relatedfacilities, material, software, services, and personnel required

m

for its operation. Examples of systems that include space

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MIL-STD-1540B (USAF)10 0CIY3BER1982

1. vehicles as a major subtier element are launch systems andon-orbit systems.

3.42 TEST DISCREPANCY

A test discrepancy is a functional or structural anomalywhich occurs during testing and which indicates a possibledeviation from specification requirements for the test item. Atest discrepancy may be a momentary, non-repeatable, orpermanent failure to respond in the predicted manner to aspecified combination of test environment and functional teststimuli. Teat discrepancies may be due to a failure of the testunit or to some other cause such as the test setup, testinstrumentation, supplied power, the test procedures, or to thecomputer software used.

3.43 TEST ENVIRONMENTS

The test environments represent conditions the unit mayexperience during ground processing and flight operations. Suchconditions normally include acoustic, vibration, acceleration,shock, thermal vacuum, nuclear radiation, and electromagneticradiation, but should also include as appropriate humidity,rain, salt spray, fungus, internal pressure, and variationa ininterface power and signal characteristics.

3.44 TEST UNIT FAILURES

A failure of the test unit is defined as a test discrepancy(see 3.42) that is due to a design, workmanship, or qualitydeficiency in the unit being tested. Non-repeatable, momentary,or any other kind of discrepancy that occurs during testing isconsidered a failure of the test unit unless it can be

I determined to have been due to some other cause.

I 3.45 THERMAL UNCERTAINTY MARGIN

A thermal uncertainty margin is included in the thermalanalysis of space vehicles to account for uncertainties inparameter such as complicated view factors, surface properties,contamination, radiation environments, joint conduction, andinadequate ground simulation. For components that have nothermal control, or have passive thermal control, tne maximumpredicted component temperatures should be at least 11 deg Cabove the maximum temperature estimated for each component based‘onmeasurements and analysis, and the minimum temperature Shouldbe at least 11 deg C below the minimum temperature estimated foreach component based on measurements and analyais. The 11 deg Cis the thermal uncertainty margin for the component. For active

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thermal control subsystems, a remaining control authority of atleast 25 percent for either or both hot and cold limits isspecified as the thermal uncertainty margin. It is used toprovide a control margin equivalent to the 11 deg C uncertaintymargin specified for passively controlled components. Forexample, if a 100 watt capacity proportional control heater isused, it should operate at 80 watts or less to maintain thecomponent above tne minimum predicted temperature. The dutycycle should be leas than 80 percent for an on-off heater. Acontrol authority margin in excess of 25 percent should bedemonstrated in cases where an 11 deg C change in theanalytically predicted component temperatures would cause thetemperature of any part of the actively controlled component toexceed an acceptable temperature limit.

3.46 ULTIMATE ~AD

The ultimate load is the maximum static load to which astructure ia designed. It is obtained by multiplying the limitload (see 3.18) by the ultimate factor of safety.

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SECTION 4

GENERAL R.EQUIRJi2iENTS

4.1 TSSTING PHIl&XOPHY

The complete test program for space vehicle hardware andsoftware encompasses development, qualification, acceptance,prelaunch validation, and other categories of testing. Testmethods, environments, and measured parameters shall be selectedto permit the collection of empirical design parameters and thecorrelation of data throughout the complete test program. Asatisfactory test program requires the completion of specifictest objectives prior to the accomplishment of others. Designsuitability is demonstrated in the development teet programprior to the start of formal qualification testing.Qualification testing should be completed prior to theinitiation of flight hardware acceptance testing.

4.1.1 Developnent Testinq. Development tests are intendedto validate hardware and computer program design concepts and toassist in the evolution of designs from the conceptual phase tothe operational phase. An objective of these tests is toidentify hardware and computer program problems early in theirdesign evolution so that any required corrective actions can betaken prior to starting formal qualification testing.Wvelopment tests may be used to confirm performance margins,manufacturability, testability, maintainability, reliability,life expectancy, and compatibility with system safety. Wherepracticable, development tests should be conducted over a rangeof operating conditions that exceed the design limits toidentify marginal design features.

4.1.2 Qualification Testing. Qualification teStSdemonstrate that adequate margine exist in the final product toassure that specification requirements are met. TO accomplishthis, the qualification test levels are usually specified as thedesign levels. These qualification test levels are establishedto exceed the range of environments and stresses expected in anysubsequent use ranging from acceptance testing through missionapplication. Por vibration testing, these margins must furtherasure that repeated acceptance testing, if necessary, will notjeopardize the integrity of the hardware. Qualification testdurations for vibration or acouatic testing of reusable hardwareshall be increased for each additional flight planned to providea sufficiently high margin to assure equipment integrity afterrepeated environmental exposures. The qualification teStS

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MIL-STD-1540B (USAF)10 OC’lY)BER1982

should validate the planned acceptance test program includingtest techniques, test procedures, test environments, groundsupport test equignnent,and computer software.

4.1.3 Acceptance Testinq. Acceptance tests are intendedto demonstrate the flight-worthiness of each deliverable item.Acceptance tests should demonstrate acceptable performance overthe specified range of mission requirements. The acceptancetests should measure performance parameters and should revealinadequacies in the manufacturing process such as workmanship ormaterial. The performance parameter measurements shouldestablish a baseline that can be used to assure that there areno data trends established in successive tests which indicate aconstant degradation of performance within specification limitsthat could result in unacceptable performance in flight.In-process production evaluation tests, burn-in or wear-intests, ,andenvironmental stress screening tests are alsoconsidered to be acceptance tests.

4.1.4 Prelaunch Validation Testinq. When a space vehicleis first delivered to the launch site, initial tests areconducted to assure space vehicle readiness for integration withthe launch vehicle. Subsequent prelaunch validation testingdemonstrates that successful integration of the space vehicle,launch vehicle, and launch facility has been accomplished, andthat compatibility exists between the space vehicle hardware andcomputer software and the launch and on-orbit systems.

4.1.5 Other Testinq. Other testing that is notdevelopment, qualification, acceptance, or prelaunch validationtesting includes on-orbit tests and testing of reuseable flighthardware.

4.1.5.1 On-Orbit Testing. Testing requirements of spacevehicles that are in orbit are an important consideration in thedesign of the space vehicle. However, the on-orbit tests are soprogram peculiar that specific requirements are not addressed inthis standard.

4.1.5.2 Reuseable Flight Hardware Testing. Reusable spacevehicle hardware consists of the space vehicles and componentsintended for repeated space missions. Airborne support-equipment and space vehicles which perform their missions whileattached to a recoverable launch vehicle are candidates forreuse, particularly for multiple mission programs. The reusableequipment would be subjected to repeated exposure to test,launch, flight, and recovery environments throughout its servicelife (see 3.37). The accumulated exposure time”of spacevehicles retained in the recoverable launch vehicle and of

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airborne support equipment is a function of the planned numberof missions involving this equipment and the retest requirementsbetween missions. Airborne support equipment environmentalexposure time is further dependent on whether or not its use isrequired during the acceptance testing of each space vehicle.In any case, the service life of reusable hardware shouldinclude all planned reuses and all planned retesting betweenuses.

The testing requirements for reusable apace hardware afterthe completion of a mission and prior to its reuse on asubsequent mission depends heavily upon the design of thereusable item and the allowable program risk. For thosereasons, specific details are not presented in this standard.Similarly, orbiting space vehicles that have completed theiruseful life spana may be retrieved by means of a recoverablelaunch vehicle, refurbished, and reused. Until some insight isprovided by experience as to the extensiveness of requiredrefurbishment, detailed test guidelines cannot be provided.Based on present approaches, it is expected that the retrievedspace vehicle would be returned to the contractor’a factory fordisassembly, physical inspection, and refurbishment. Alloriginally specified acceptance tests should be conducted beforereuse.

4.2 Tt%T PLANS AND PROCEDURES

4.2.1 Preparation. The contractor shall establishprocedures for performing all required tests in accordance withdetailed test plans approved by the contracting officer. Thetest plans shall be based upon a function by function missionanalysis and the test requirements. The teat plana shouldindicate the test requirements~ testin9 approach for each item,related special test equipment, facilities, and SYStem interfacerequirements. The test plans should identify the allocation ofrequirements to appropriate testable levels of aasembly.Traceability shall be provided from the specified requirementsto the test-procedures; The test procedures shall cover alloperations in enough detail so that there is no doubt as to whatis to be done. The pass-fail test criteria shall be determinedprior to the start of every test.

4.2.2 Test Condition lblerances. The test conditiontolerances allowed by this standard shall be applied to thenominal test values specified. Unless otherwise specified, thefollowing maximum allowable tolerances on test conditions shallapply.

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Temperature ~3deg C

PressureAbove 1.3 x 102 pascal1.3 Xlo-1 to 1.3 Xlo 2 :a:c::r)

(0.001 Torr to 1 Torr)wss than 1.3 x 10-1 pascals

( 0.001 Torr)

Relative Humidity

Acceleration

Vibration Frequency

Sinusoidal Vibration Amplitude

~ 10 percent~ 25 percent

~ 80 percent

~ 5 percent

~ 10 percent

~ 2 percent

~ 10 percent

Random vibration AccelerationPower Spectral Density

20 to 500 Hz (25 Hz or narrower) + 1.5 dB500 to 2000 Hz (50 Hz or narrower) = 3.0 dB

Random Overall grins ~ 1.5 dB

Sound Pressure1/3 OctaveOverall

Shock Response1/6 OctaveAmplitude

Static bad

LevelBand + 3.0 dB

~ 1.5 dB

Spectrum (Q = 10)Band Center Frequency ~ 6 dB with 30

percent of theresponsespectrum centerfrequencyamplitudesgreater thannominal testspecification

~ 5 percent

4.2.3 Tailorinq. This standard specifies testrequirements that have been shown to assure high reliability inspace missions. However, it is intended that these testbaselines should be tailored to each space program consideringdesian comDlexitv. state of the art. mission criticality. cost,and ;ccept;ble risk. Forrelax the requirements inprograms the requirements

some space programs this tail~iing maythis standard, while for othermay be made more stringent to reduce

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MIL-STD-1540B (USAF)10 OCIXIBER 1982

risk of on-orbit failures or to demonstrate with greaterconfidence that the space vehicle or components performadequately when all parameters, environments, and relateduncertainties are considered. For example, the optional testsshown in the test baseline tables should be added as requiredtests, where appropriate, as determined from considerations ofdesign features, required lifetime, environmental exposure, andexpected useage. The tailoring is a continuing processthroughout the acquisition that should be implemented by thewording used to state the testing requirements in thespecifications of the space system, space vehicles, andcomponents or in other applicable contractual documents.

4.3 RETEST

If a test discrepancy (see 3.42) occurs, the test should beinterrupted and the discrepancy verified. If the discrepancy isdispositioned as due to the test setup, software, or to afailure in the test equipment, the test being conducted at thetime of the failure may be continued after the repairs arecompleted, as long as the discrepancy did not result in anoverstress test condition. If the discrepancy is dispositionedas a failure in the item under test, the preliminary failureanalysis and appropriate corrective action shall normally becompleted before testing is resumed. If the failure occursduring system testing, the test may be continued if thediscrepant area is not affected by the continuation of testing.

The conducting of a proper failure analysis plays animportant part in the decision on the type of retest. It shouldinclude the determination of whether a failure occurred, thecause of the failure, the physics of the failure, and isolationof the failure to the smallest replaceable item. The degree ofretest shall be determined for each case based upon the natureof the failure. In the case of a significant redesign of acomponent, all previous qualification testa shall be repeated.After significant component rework, all previous acceptancetests except burn-in shall be repeated. In the case ofextensive component rework, repetition of the burn-in is alsorequired. Where the redesign or rework of the component.is veryminor, it may be acceptable to only repeat functional testingand the test in which the failure occurred.

Where significant redesign or rework of components isrequired as the result of failure during system level testing,the system level test in which the failure occurred, aa well asfunctional testing of the failed subsystem, shall be repeated.Repetition of system level environmental tests may be necessaryif the redesign was extensive or the number of components

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MIL-STD-1540B (USAF)10 OC’IX3BER1982

●changed out and connectors demated is so large as to reduceconfidence in the space vehicle.

4.4 TEST DATA ANALYSIS

A test data bank containing all pertinent system, spacevehicle, subsystem, and component test data taken throughout theprogram shall be maintained. To permit as complete anevaluation as possible of component, subsystem, and spacevehicle performance under the various specified test conditions,all relevant test measurements and the environmental conditionsimposed on the units shall be recorded on magnetic tape or byother suitable means. These records are intended for post-testanalysis to supplement the real-time monitoring and tofacilitate the mechanized accumulation of trend data for thecritical test parameters. Test data shall be examined for outof tolerance values and for characteristic signatures.Transient responses and mode switching tests shall be examinedfor proper response. The test data shall also be comparedacross major test sequences for trends or evidence of anomalousbehavior.

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SECTION 5

DEVE~PMENT TESTING

5.1 GENERAL

The objective of the develoganent tests is to assure tnattesting of critical items at all levels of assembly issufficient to validate the design approach. Requirements fordevelopment testing therefore depend upon the maturity of thesubsystems and components used and upon the operationalrequirements of the specific program. Development tests arenecessary to validate new design concepts and the application ofproven concepts and techniques to a new configuration.Development tests are also conducted to verify design criteriafor structures and components and to determine design marginsand failure modes. llevelo~ent tests may be conducted onbreadboard equipment, prototype hardware, or the developmenttest vehicle equipment and software. when development tests areproposed on qualification or flight hardware, the approval ofthe contracting officer is required.

By its nature, development testing cannot be reduced to astandardized set of procedures. The development testrequirements are necessarily unique to each new space vehicle.It is not the intent of this section to define the requireddevelopment tests, but to provide guidelines for conductingappropriate tests when their need has been established.

5.2 CQMPONENT DEVEU)Pt4ENT TESTS

The major portions of the development tests are conductedon breadboards and prototypes at the subassembly and componentlevels. The object;ve i.s-;arlyverification of-the criticaldesign concepts to reduce the risk involved in committing thedesign to qualification and flight hardware. The emphasis forhardware is on packaging design, electronic and mechanicalperformance, and capability to withstand environmental stress.New designs should be characterized across worst case voltage,frequency, and temperature variations at the breadboard level.Functional testing in thermal and vibration environments isnormally conducted. Development tests of deployable and of theattitude control subsystem are normally conducted. Life testsof critical itemsmoving mechanicalconducted.

whi~h may have a wea~out failure mode, such asassemblies (see 3.26), should also be

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For electronic boxes, thermal mapping in a vacuumenvironment for known boundary conditions may be needed toverify the internal component thermal analysis. The correlatedthermal model is then used to demonstrate that critical piecepart temperature limits, consistent with reliabilityrequirements and performance, are not exceeded. When electronicbox packaging is not accomplished in accordance with known andaccepted techniques relative to the interconnect system, partsmounting, board sizes and thickness, number of layers, thermalcoefficients of expansion, or installation method, developmenttests should be performed. The tests should establishconfidence in the design and manufacturing processes used.Temperature cycling and random vibration testing should be

I conducted to evaluate the entire package.

5.3 SPACE VEHICLE AND SUBSYSTEM DEVE~PMENT TESTS

Tests of structural and thermal development models areI

I often necessary to confirm dynamic and thermal environmentalcriteria for design of subsystems, to verify mechanical

I interfaces, and to assess functional performance of deploymentmechanisms and thermal control systems. Space vehicledevelopment testing also provides an opportunity to develophandling and operating procedures as well as to understandsystem interactions.

For recoverable launch vehicles, a mechanical fit andoperational interface test with the facilities at the launch-.Site is recommended. The flight structure should be used ifpossible; however, a facsimile or portions thereof may be used.toconduct the development tests at an early point in theschedule in order to reduce the impact of hardware designchanges that may be necessary.

A modal survey is normally conducted to define or verify ananalytically derived dynamic model of the space vehicle for usein launch vehicle flight loading event simulations and for usein examinations of post-boost configuration elastic effects uponcontrol precision and stability. This test is conducted on aflight quality structural subsystem as augmented by masssimulated components. The data obtained should be adequate todefine orthogonal mode shapes, mode frequencies, and modedamping ratios of all modes which occur within the frequencyrange of interest. In most instances, modes in the frequencyran9e from zero to 50 Hz should be measured.

A structural development test is also needed to verify thestiffness properties of the space vehicle and to measure memberload and stress distributions and deflections in structures with

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redundant load patha. The stiffness data are of particularinterest where nonlinear structural behavior exist that is notfully exercised in the modal survey. This may include nonlinearbearings, elastic buckling of panels, gapping at preloadedinterfaces, and slipping at friction joints. The member loadand stress distribution data are used to experimentally verifythe Loads Transformation Matrix (L’IM). Deflection data are alsoused to experimentally verify the Deflection TransformmationMatrix (D’IM). These matrices are used, in conjunction with thedynamic model, to calculate member loads (axial forces, bendingmoments, shears, torsional moments), and various stresses anddeflections, which are converted into design and clearancemargins for the space vehicle. This development test could bedone in conjunction with the modal survey on the complete spacevehicle or, alternatively, it could be done at a subcontractor’sfacility where significantly large subassemblies could beverified prior to shipping to the integrating contractor. Thisdevelopment test does not replace the structural static loadtest (see 6.3.1) that is required for vehicle qualification:however, the two tests may be incorporated into a single testsequence that encompasses the requirements of both tests.

Since high frequency vibration and shock responses acedifficult to predict by analytical techniques, acoustic andshock development testing of the space vehicle may be necessaryto verify the adequacy of the dynamic design criteria forcomponents of the space vehicle. Space vehicle components thatare not installed at the time of the test should be dynamicallysimulated as closely as practical with respect to maas, centerof gravity, moments of inertia, interface stiffness, andgeometric characteristics. For the acoustic test the spacevehicle is normally exposed to the qualification environmentan acoustic ChSMber. Since the acoustic test is intended toexcite a random vibration environment throughout the spacevehicle which simulates the environment experienced durina

in

launch and flight, a test with the vehicle-enclosure inst~lled“is highly desirable for vehicles that have an enclosure as apart of the flight configuration. For the pyro shock test allpyrotechnically operated devices and other equipment capable ofimparting a significant shock impulse to the space vehicleshould be operated. When significant shock levels are expectedfrom subsystems not on board the space vehicle under test, suchas the launch vehicle separation shock, the adaptor subsystem orsuitable simulation shall be attached and appropriatepyrotechnics or other means used to simulate the shock imposed.Since the pyro shock environment may vary significantly betweentest exposures, a minimum of three tests of each dominantpyre-device should be used to simulate the maximum predictedshock environment (see 3.22).

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MIL-STD-1540B (USAF)10 OC’lTIBER1982

A thermal balance development test may be necessary toverify the analytical modeling of the space vehicle andcom~nent thermal design criteria. For space vehicles in whichstructural thermal distortions are critical to mission success,this test also evaluates potential alignment problems. Theconfiguration of the space vehicle should consist of a thermallyequivalent structure with addition of equipment panels, thermalcontrol insulation, finishes, and thermally equivalent models ofelectronic, pneumatic, and mechanical components. ‘restingisconducted in a space simulation test chamber capable ofsimulating the ascent, transfer orbit, and orbital thermalvacuum conditions.

The dynamic environment experienced by the space vehicleduring road transportation is normally controlled to levels lessthan the maximum levels predicted for launch and flight. Manalysis of space vehicle response to the transportationenvironment requires definition of the road surface. Since thelatter is difficult to define, it is often necessary to conducta development test of the space vehicle and transporter over arepresentative course to’verify that the space vehicle would notbe damaged. For such a test, a space vehicle development modelor a simulator which has space vehicle mass properties.may beused, and instrumentation would be installed to measure bothspace vehicle and transporter dynamic responses.

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SECTION 6

QUALIFICATION TESTING

GENERAL QUALIFICATION TEST REQUIREMENTS

6.1.1 @alification Hardware. The space vehicle hardwarefor qualification shall be produced from the same drawings,

using the-same materials, toolin~, and methods as flighthardware, except as modified to accommodate minor changes thatmay be necessary to conduct the test. These minor changes mayinclude adding such items as thermocouples, strain gages, andmonitoring leads. The qualification hardware shall havedemonstrated successful performance at acceptance levels inacoustic, vibration, and thermal environments prior toproceeding to the higher levels for formal qualification testingin that environment. For those environments which are appliedby axis, both the acceptance and qualification tests can becompleted in one axis before switching to another.

6.1.2 f)ualification Test Levels. The qualification testlevels shall at least equal the design levels. If the equipmentis to be used by more ti-anone program orvehicle locations, the qualification testthe design levels of the various programslocations involved.

6.2 SPACE VEHICLE QUALIFICATION TESTS

The space vehicle qualification testall the required tests specified in Table

in different spacelevels shall envelopeor space vehicle

baseline consists of1. The test baseline

shall be tailored for each program, giving consideration to boththe required and optional tests; however, deviations from the

I baseline requirements for the required tests shall be approvedby the contracting officer. Additional special tests such asalignments, instrument calibrations, antenna patterns, and masarwoperties that are conducted as acceptance tests for flight;ehicles shall be conducted on the qu~lification flight v=hicleunit. If the space vehicle is controlled by on-board dataprocessing, the flight version of the computer software shall beresident in the space vehicle computer for these tests. Theverification of the operational requirements shall bedemonstrated to the extent practicable.

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Table I. Space Vehicle Qualification Tests

Reference Suggested Required (R)Test Paragraph Sequence OR

Optional (0)

Functional 6.2.1 ~(1) R

EMc 6.2.2 2 R

Acoustic 6.2.3 5 R(2)

Vibration 6.2.4 5 0

Pyro Shock 6.2.5 4 R

Pressure 6.2.6 3, 6 R

Thermal Vacuum 6.2.7 9 R

Thermal Balance 6.2.8 8 R

Thermal ~cling 6.2.9 7 0(3)

Notes: (1) Electrical and mechanical functional tests shallbe conducted prior to and following eachenvironmental test.

(2) Conduct vibration in place of acoustic test forvehicles of compact shape and with weight lessthan 180 kilograms.

(3) Required if thermal cycling acceptance test 7.1.8is conducted.

6.2.1 Functional Test, Space Vehicle Qualification.

6.2.1.1 Purpose. This test verifies that mechanical andelectrical performance of the space vehicle meets thespecification requirements, ,verifies compatibility with groundsupport equipment, and validates all test techniques andsoftware algorithms used in computer-assisted commanding anddata processing.

6.2.1.2 Mechanical Functional Test. Mechanical devices,valves, deployable, and separable entities shall befunctionally tested, with the space vehicle in the launch,orbital, or recovery configuration appropriate to the function.Alignment checks shall be made where appropriate. The maximum

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and minimum limits of acceptable performance shall be determinedwith respect to mechanics, time, and other applicablerequirements. For each mechanical operation, such as appendagedeployments, tests shall demonstrate positive margins ofstrength, torque margins, and that they function at conditionsabove and below specified operational limits. Where operationin a l-g environment cannot be performed, a suitable ground testfixture may be utilized to permit operation and evaluation ofthe devices. Fit checks shall be made of the space vehiclephysical interfaces with the launch vehicle by means of launchvehicle master gages or interface assemblies.

6.2.1.3 Electrical Functional Test. The space vehicleshall be in its fliqht configuration with all components andsubsystems connected except ~yrotechnic elements.- This testshall verify the integrity Of all electrical circuits in thespace vehicle, including redundant paths, by the application ofan initiating stimulus and the confirmation of the successfulcompletion of the event. The test shall be designed to operateall components, primary and redundant, and all commands shall beexercised. The operation of all thermally controlledcomponents, such as heaters and thermostats, shall be verifiedby test. The test shall demonstrate that all commands havingpreconditioning requirements (such as enable, disable, specificequipment configuration, specific command sequence, etc.) cannotbe executed unless the preconditions are satisfied. Whereverpossible, equipment performance parameters (such as power,voltage, gain, frequency, command and data rates) shall bevaried over specification ranges to demonstrate the performancemargins. Autonomous functions shall be verified to occur whenthe conditions exist for which they are designed. The spacevehicle main bus shall be continuously monitored by a powertransient monitor system. All telemetry monitors shall beverified, and pyrotechnic circuits shall be energized andmonitored. A segment of this test shall operate the spacevehicle through a mission profile with all events occurring inactual flight sequence to the extent practical. This sequenceshall include the final countdown, launch, ascent, separationfrom the launch vehicle, orbital injection, orbital o~eration,and return from orbit as appropriate.

6.2.1.4 Supplementary Requirements. Mechanical andelectrical functional tests shall be conducted prior to andafter each of the environmental tests to detect equiganentanomalies and to assure that performance meets the requirements .of the specification. These tests do not require the missionprofile sequence. Sufficient data analysis to verify theadequacy of the testing and the validity of the data shall becompleted before disconnection from a particular environmental

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MIL-STD-1540B (uSAP)10 OC’lY3BER1982

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test configuration in order that any retesting required can bereadily accomplished. During these tests, the maximum use oftelemetry shall be employed for data acquisition, problemidentification, and problem isolation.

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6.2.2 Electromagnetic Compatibility Test, Space VehicleQualification

6.2.2.1 Purpose. This test demonstrates electromagneticcompatibility of the space vehicle and ensures the space vehiclehas adequate margins in a simulated launch, orbital, and returnfrom orbit electromagnetic environment.

6.2.2.2 Test Description. The operation of the spacevehicle and selection of instrumentation shall be orientedtoward determining the margin against malfunctions andunacceptable or undesired responses due to electromagneticincompatibilities. The test shall demonstrate satisfactoryelectrical and electronic equipment operation in conjunctionwith the expected electromagnetic radiation from other systemsor equipment such as the launch vehicle and the ground supportequipment. The space vehicle shall be subjected to the requiredtests while in the launch, orbital, and return from orbitconfigurations and in all possible operational modes. Specialattention shall be given to areas indicated to be marginal byanalysis. Potential electromagnetic interference from the spacevehicle to other systems shall be measured. The tests shall be ●conducted in accordance witn the requirements of MIL-STD-1541.Electroexplosive devices with bridge wires, but otherwise inert,shall be installed in the system and monitored during all tests.

6.2.3 Acoustic Test, Space Vehicle Qualification

6.2.3.1 Purpose. This test demonstrates the ability ofthe space vehicle to withstand or, if appropriate, to operate inthe design level acoustic environment which is the maximum levelimposed in flight plus a design margin. This test also verifiesthe adequacy of component vibration qualification criteria.

6.2.3.2 Test Description. The space vehicle shall beinstalled in a reverberant acoustic cell capable of generatingdesired sound pressure levels. It shall be mounted on aflight-type support structure or reasonable simulation thereof.The mechanical configuration of the space vehicle shall be as itis during ascent (for example, solar arrays and antennasstowed) . Where possible, qround handlinq equiDment and testequipment shall heshall be installedpoints of critical

removed. Adequate dy~am~c ~nstrumentationto measure vibration responses at attachmentand representative components.

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6.2.3.3 Test Levels and Duration. The acoustic testspectrum shall be the design environment (see 3.8) which is themaximum predicted flight environment (see 3.20) PIUS the desi9nmargin (6 dB : see 3.12). However, the overall sound pressucelevel of the qualification test shall not be less than 144 dB.Exposure test time shall be at least three times the expectedflight exposure time to the maximum flight environment, or threetimes the acceptance test duration if that is greater, but notless than 3 minutes. Operating time should be dividedapproximately equally between redundant circuits. Whereinsufficient time is available at the full test level to testall redundant circuits, all functions, and all modes, extendedtesting at a level 6 dB lower shall be conducted as necessary tocomplete functional testing.

6.2.3.4 Supplementary Requirements. During the test allelectrical and electronic components, even if not operatingduring launch, shall be elect~ically energized and sequencedthrough operational modes to the maximum extent possible withthe exception of components that may sustain damage ifenergized. Continuous monitoring of several perceptiveparameters shall be provided to detect intermittent failures.Functional tests are required before and after the environmentalexposure.

6.2.4 Vibration 2&st, Space Vehicle Qualification

6.2.4.1 Purpose. This test demonstrates the ability of aspace vehicle to w~thstand or, if appropriate, to opecate in thedesign level vibration environment which is the maximum levelimposed in flight plus a design margin. ThlS test alao verifiesthe adequacy of component vibration qualification criteria.This test is only used for small vehicles with weights under 180kg and with compact shapes.

6.2.4.2 Test Description. The space vehicle shall beattached to a vibration fixture by a flight type adapter.ViOration shall be applied at the-base of the adapter via thefixture in each of three orthogonal directions, one directionbeing parallel to the thrust axis. Generally, antennas, booms,and solar arrays shall be stowed in the launch condition.Fdequate dynamic instrumentation shall be installed to measurevibration responses in 3 axes at attachment points of criticaland representative components.

6.2.4.3 Test Levels and Duration. The test levels shallbe chosen to produce vibration responses within the equipmentwhich are the maximum predicted flight environment (see 3.24)plus the design margin (6 dB : see 3.12). Structural reseonses

33



Ishall be limited at resonant frequencies so that structural

1loads do not exceed design limit-loads (see 3.18). Testduration in each axis shall be three times the expected flightexposure time to the maximum flight environment, or three timesthe acceptance test duration if that is greater, but not lessthan 3 minutes. Operating time should be divided approximately

I equally between redundant circuits. Where insufficient time isavailable at the full test level to test all redundant circuits,all”functions, and all modes, extended testing at a level 6 dBlower shall be conducted as necessary to complete functionaltesting.

I

6.2.4.4 .!iupplementary Requirements. For small vehicleswith weights of under 180 kilograms (kg) and compact shapes, avibration test shall be conducted in lieu of an acoustic test.In these cases, special care is required to ensure that loads inexcess of design loads are not imposed on the space vehicle.Items such as resonance effects of the structure, adapter type,table control techniques, and location of controlaccelerometers, shall be carefully considered when vibration isselected in lieu of acoustic testing. During the test, allelectrical and electronic components, even if not operatingduring launch, shall be electrically energized and sequencedthrough operational modes to the maximum extent possible (withthe exception of components that may sustain damage if-energized). Continuous monitoring of several perceptiveparameters shall be provided to detect intermittent failures.Functional tests are required before and after the environmentalexposure. .

6.2.5 Pyro Shock Test, space Vehicle Qualification

I 6.2.5.1 Purpose. This test demonstrates the capability ofthe space vehicle to withstand or, if appropriate, to operate inthe design level pyro shock environments which are the levelspredicted for flight plus a design margin. This test also

I verifies the adequacy of component pyro shock criteria.

6.2.5.2 Test Description. In this test or series of testsegments, all pyrotechnically operated devices and otherequipment capable of imparting a significant shock impulse tothe space vehicle shall be operated. Separation subsystemshocks are often more severe than those from other pyrotechnicdevices, and operation of the separation subsystems is thereforeparticularly significant. For these tests, the space vehicleshall be suspended or otherwise supported so as to preclude thepossibility of recontact between separated portions thereof.When significant shock levels are predicted from subsystems noton board the space vehicle under test, such as the launch

34



vehicle separation shock, the adapter subsystem or suitablesimulation shall be attached and appropriate pyrotechnics orother means used to simulate the shock imposed. Adequatedynamic instrumentation shall be installed to measure pyro shockresponses in 3 axes at attachment points of critical andrepresentative components.

Support of the space vehicle varies with the configurationand may vary during the course of this test series. ‘lbpermitoptimum positioning and prevent damage to such items asdeployment booms, paddles, and ejectable, a series ofindividual test setups or deployment restraints may berequired. The test setup shall permit, as nearly as possible,flightlike dynamic response of the space vehicle structure.

I

I

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6.2.5.3 Test Levels and Duration. All pyrotechnic devices(e.g., explosive bolt, nut, pin puller, marmon clamp, etc.)shall be fired at least one time. Those pyrotechnic devicesproducing shock levels within 6 dB of the maximum shock responsemeasured from any of the devices shall be fired two additionaltimes to provide the expected variability in the shockenvironment. Firing Of both primary and redundant pyres shallbe in the same sequence as they are designed to fire in flight.

6.2.5.4 Supplementary Requirements. Electrical andelectronic components shall be operating and monitored to themaximum extent possible. Functional tests are required beforeand after environmental exposure.

6.2.6 Pressure Test, Space Vehicle Qualification

6.2.6.1 Purpose. This test demonstrates the capability offluid subsystems to meet the flow, pressure, and leakage raterequirements specified.

6.2.6.2 Test Description. The space vehicle shall beplaced in a facility that provides the services and safetyconditions required to protect personnel and equipment duringthe testing of high pressure subsystems and in the handling ofdangerous fluids. ‘R?stsshall be performed to verifycompatibility with the test setup and to ensure that propercontrol of the equipment and test functions is provided. Therequirements of the subsystem including flow, leakage, andregulation shall be measured while operating applicable valves,pumps, and motors. The flow checks shall verify that theolumbina confiauratians are adeauate. Checks for subsystem~leanli~ess, m~isture levels, aid PH shall alsopressurized subsystems are assembled with otherwelded connections, the specified torque values

be made. Wherethan brazed orfor these

35



I

I

connections shall be verified prior to leak checks.

In addition to the high pressure test, propellant tanks andthruster valves shall be tested for leakage under propellantservicing conditions. The system shall be evacuated to theinternal pressure normally used for propellant loading and thesystems pressure monitored for any indication of leakage.

6.2.6.3 Test Levels and Duration. The subsystem shall bepressurized to proof pressure (see 3.34) and held for 5 minutes,then the pressure shall be reduced to the maximum predictedoperating pressure (see 3.21). Unless specified otherwise, theproof pressure equals 1.5 times the maximum operating pressure.This sequence shall be conducted three times. Inspection forleakage after these cycles shall be at the maximum operatingpressure. The duration of the evacuated propulsion system leaktest shall not exceed the time that this condition is normallyexperienced during propellant loading.

6.2.6.4 SUpp lementary Requirements. Applicable safetystandards shall be followed in conducting all tests. Speciallyformulated bubbleforming solutions are suitable for detectingexternal leakage at such locations as joints, fittings, plugs,and lines, where the allowable limits are from 0.00001 to 0.01cubic centimeters per second (cubic cm/see). Solutions that areused for detecting leaks shall be compatible with the media ●being leak tested or with the media which could contact anyresidues. Liquid displacement methods may be used for detectingleakage through poppet seats and internal seals for measurementrequirements of 0.1 to 30 cubic cm/sec. Helium or radioactivetracer gas leak detectors may be used for leakage rates from0.0000001 to 0.0001 cubic cm/sec. The use of halogen gasdetectors for liquid propulsion subsystems shall be avoided.Leak tests shall be conducted only after satisfactory proofpressure tests have been completed. Leak detection andmeasurement procedures may require vacuum chambers, bagging ofthe entire space vehicle or localized areas, or other specialtechniques to achieve the required accuracies.

6.2.7 Thermal Vacuum Test, Space Vehicle Qualification

6.2.7.1 Purpose. This test demonstrates the ability ofthe space vehicle to meet design requirements under vacuumconditions and temperature extremes which simulate thosepredicted for flight plus a design margin.

6.2.7.2 Test Description. The space vehicle shall beplaced in a thermal vacuum chamber and a functional testperformed to assure readiness for chamber closure. The vehicle

●36


MIL-STD-1540B (USAF)

I10 OCTOBER 1982

1 shall be zoned into separate equipment areas based on thelocation of critical components within each area. Componentsthat operate during ascent shall be monitored for corona, andmultipacting (see 3.27) as applicable, as the pressure isreduced to the lowest specified level. Equipment that does notoperate during launch shall have electrical power applied afterthe test pressure level has been reached. A temperature cycle

,. begins with the space vehicle at ambient temperature. Thetemperature is reduced to the specified low level and

1- stabilized. Component temperature stabilization has beenachieved when the rate of temperature change is no more than 3deg C/hour. Following the cold soak, the temperature shall beraised to the highest specified level and stabilized. Followingthe high temperature soak, the space vehicle shall be returnedto ambient temperatures to complete one temperature cycle.Functional tests shall be conducted during the first and last

Itemperature cycle at both the high and low temperature limitswith functional operation and monitoring of perceptiveparameters during all other cycles. In addition to thetemperature cycles, the chamber shall be programmed throughvarious orbital operations. Operational sequences shall becoordinated with expected orbital environments, and a completecycling of all equipment shall be performed including the

e

operating and monitoring of redundant equipment and paths.System electrical equipment shall be operating and monitoredthroughout the test. Strategically placed temperature monitorsshall assure attainment of temperature limits. Strategicallyplaced witness plates and quartz crystal microbalances or otherinstrumentation shall be installed in the test chamber to assurethat outgassing from the space vehicle and test equipment doesnot degrade system performance beyond specified limits.

6.2.7.3 Test Levels and Duration. Temperatures in variousequipment areas shall be controlled by the external testenvironment and internal heating resulting from equipsentoperation so that during the hot cycle the temperature on atleast one component in each equipnent area at its design hightemperature and one component dur’ing the cold cycle is at itsdesign low temperature. The temperature extremes shall beestablished by a survey of predicted temperatures in variousequi~ent areas and may have to be adjusted to the performanceof the most sensitive components in a particular area.Temperatures on the components shall not be allowed to exceedthe design levels for the components. The pressure shall bemaintained at 0.0133 pascals (0.0001 ‘lbrr)or less. All orbitaloperational conditions and all equipment functional modesincluding redundancy shall be tested. The qualification testshall include at least eight complete hot-cold cycles at themaximum predicted orbital rate of temperature change and with at

37



least an 8-hour soak at each temperature extreme. Operatingtime should be divided approximately equally between redundantcircuits.

6.2.7.4 Supplementary Requirements. Since the purpose ofthe more severe temperature extreme IS to demonstrate anadequate design margin, it may be necessary to force temperatureextremes at certain locations by altering thermal boundaryconditions locally or by altering the operational sequence toprovide additional heating or cooling. Adjacent equipments maybe turned on or off; however, any special conditioning withinthe space vehicle shall generally be avoided. Externalbaffling, shadowing, or heating shall be utilized to the extentfeasible.

6.2.8 Thermal Balance Test, Space Vehicle Qualification

6.2.8.1 Purpose. This test verifies the analyticalthermal model and demonstrates the ability of the space vehiclethermal control subsystem to maintain components, subsystems,and the entire space vehicle within the specified operationaltemperature limits. This test also verifies the adequacy ofcomponent thermal desiqn criteria.

6.2.8.2 Test Description. The qualification space vehicleshall be tested to simulate the thermal environment seen by the ●space vehicle during the transfer orbit and orbital missionphases. Tests shall be conducted over the full mission range ofseasons, equipment duty cycles, solar angles, and eclipsecombinations so as to include the worst case high and lowtemperature extremes for all space vehicle components. Specialemphasis shall be placed on defining the test conditionsexpected to produce the maximum and minimum batterytemperatures. Sufficient measurements shall be made on thespace vehicle internal and external components’ to effectverification of the space vehicle thermal design and analyses.The power requirements of all thermostatically controlledheaters shall be verified during the test. The test chamber,with the test item installed, shall provide a pressure of 0.0133pascals (0.0001 Torr), or less. Where appropriate, provisionsshall be made to.prevent the,test item from “seeing” warmchamber walls by using black-coated cryogenic shrouds of.sufficient area and shape that are capable of approximatingliquid nitrogen temperatures. The space vehicle thermalenvironment may be supplied by one of the following three

I methods:

a. Method I. Absorbed Flux. The absorbed solar,albedo, and planetary irradiation is simulated

38


●


using heater panels or IR spectrum adjusted forthe external thermal coating properties andprojected by IR lamps or heater panels.

b. Method II. Incident Flux. The intensity,spectral content, and angular distribution of theincident solar, albedo, and planetary irradiationis simulated.

c. Method III. Combination. Thermal environment is1- supplied by a combination of incident and

absorbed irradiation.

I The selection of the method and fidelity of the simulationdependa u~n detaila of the space vehicle thermal design such asvehicle geometry, the size of internally produced heat loadscompared with those supplied by the external environment, and

I the thermal characteristics of the externaL surfaces.Instrumentation shall be incorporated down to the componentlevel to evaluate total space vehicle performance withinoperational limits as well as to identify component problems.The space vehicle shall be operated and monitored throughout theteat. Dynamic orbital simulation of the space vehicle thermal

o

environment shall be provided unless the external space vehicletemperature doea not vary significantly with time. For example,static simulation is usually adeq”uate for spinning spacevehicles.

6.2.8.3 Test Levels and Duration. Test conditions anddurations for this test are dependent upon the space vehicleconfiguration, design, and mission details. Normally, boundaryconditions for evaluating thermal design shall include: (a)maximum external absorbed flux plus maximum internal

1- dissipation, and (b) minimum external absorbed flux Plus minimuminternal power dissipation. The thermal time constant of thesubsystems and orbital maneuvering both influence the timerequired for the space vehicle to achieve thermal equilibriumand hence the test duration. Thermal equilibrium has beenachieved when the equipnent temperature change is no more than 3deg C/hour. The tests should simulate the full range ofseasona, equifanentduty cycles, solar angles, and eclipsecombinations so as to produce the worst caae high and lowtemperature extremes for all space vehicle components.

6.2.8.4 Supplementary Requirements. This test augmentsand validates the detailed thermal analysis. Pass criteriadepend not only on survival and operation of each equipmentwithin sDecified tem~rature limits, but also on correlation of

‘6the test-with theore~ical thermal models. As a goal,

39


MIL-STD-1540B (uSAP)10,OC~BER 1982

COKKelatiOn Of test results to the thermal model predictionsshould be within ~ 3 deg C. Lack of correlation with thetheoretical models may indicate either a deficiency in themodel, test setup, or space vehicle hardware. The thermalbalance test can be combined with the thermal vacuum test. Thecorrelated thermal math model is then used to make the finaltemperature predictions for the various mission phases, e.g.,prelaunch, ascent, on-orbit, etc. The thermal margins are thenbased on these final temperature predictions.

6.2.9 Thermal Cycling Test, Space Vehicle Qualification

6.2.9.1 Purpose. ThlS test demonstrates the ability ofthe space vehicle to withstand the thermal stressing environmentof the space vehicle thermal cycling acceptance test (7.1.8)plus a design margin.

I6.2.9.2 Test Description. The space vehicle shall be

placed in a thermal chamber at ambient pressure, and afunctional test shall be performed to assure readiness for thetest. The space vehicle shall be operated and monitored duringthe entire test, except that space vehicle power may be turnedoff if necessary to reach stabilization at the coldtemperature. Space vehicle operation shall be asynchronous withthe temperature cycling, and redundant circuits shall beoperated with approximately equal time on each redundant ●circuit. Unfavorable combinations of temperature and humidityshall be avoided so there is no moisture deposition either onthe exterior surfaces of the space vehicle or inside spaceswhere the humidity is slow to diffuse, e.g., multilayezinsulation and enclosed electronic equipment. When the relativehumidity of the inside spaces of the space vehicle is below thevalue at which the cold test temperature would causecondensation, the temperature cycling shall begin. One completetemperature cycle is a period beginning at ambient temperaturethen cycling to one temperature extreme and stabilizing, then tothe other temperature extreme and stabilizing, and thenreturning to ambient temperature. Strategically placedtemperature monitors installed on components shall assureattainment and stabilization of the temperature extremes atseveral components. Auxiliary heating and cooling may beemployed for selected temperature-sensitive components, e.g.,batteries. If it is necessary to achieve the temperature rateof change, parts of the space vehicle such as solar panels andpassive thermal equipment may be removed for the test. The lasttemperature cycle shall be a soak cycle during which the spacevehicle shall remain at each temperature extreme while afunctional test, including testing of redundant circuits, isconducted.

40



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6.2.9.3 Test Levels and Duration. The space vehicletemperature range from hot to cold shall be the maximum possiblewithin the constraints of the component design temperatures.The minimum space vehicle temperature range should be 70 deg C.Auxiliary heating and cooling may be used to protect selectedtemperature sensitive components. The average rate of change oftemperature from one extreme to the other shall be as rapid aspossible. The test shall include 25 percent more thermal cyclesthan the thermal cycling acceptance t@St (7.1.8).

6.3 SUBSYSTEM QUALIFICATION TESTS

Subsystem level qualification tests shall be conducted onthose subsystems that are subjected to environmental acceptancetests. Qualification tests shall also be conducted at thesubsystem level where this level of testing provides a morerealistic test simulation, such as the required structuralstatic load test. At the other extreme, the qualification ofcertain components such ss interconnect tubing or wiring may bemost readily completed at the subsystem level of assembly ratherthan at the component level. In this case, the subsystem may betreated as a component, and the appropriate component leveltests conducted at the subsystem level to complete requiredcomponent qualification tests.

6.3.1 Structural Static Load Test, Subsystem Qualification

6.3.1.1 Purpose. This test demonstrates the adequacy ofthe structure to meet requirements of strength or stiffnesS, orboth, with the desired design margin when subjected to simulatedcritical environments, such as temperature and loads, predictedto occur during its service life.

6.3.1.2 Test Description. The structural configuration,materials, and manufacturing processes employed in the

I qualification test specimens shall be identical to those offlight articles. When structural, items are rebuilt orreinforced to meet soecific strendth or riaidity requirements.

I

.e

all modifications sh>ll be structfirally id~ntic~l t6 the changesincorporated in flight articles. The support and loadapplication fixture shall consist of an adequate replication ofthe ad]acent structural section to provide boundary conditionswhich simulate those existing in the flight article. Staticloads representing the design limit load and the design ultimateload [see 3.46) shall be applied to the structure, andmeasurements of the strain and deformation shall be recorded.Strain and deformation shall be measured before loading, afterremoval of the limit loads, and at several intermediate levelaup to limit load for post-teat diagnostic purposes. The test


MIL-STD-1540B (USAF)10 OCTt)BER 1982

conditions shall include the combined effects of acceleration,pressure, preloads, and temperature. These effects can besimulated in the test conditions as long as the failure modesand design margins are enveloped by the simulations. Forexample, temperature effects, such as material degradation andadditive thermal stresses, can often be accounted for byincreasing mechanical loads. Analysis of flight profiles shallbe used to determine the proper sequencing or simultaneity forapplication of thermal stresses. When prior loading historiesaffect the structural adequacy of the test article, these shallbe included in the test requirements. If more than one ultimateload condition is to be applied to the same test specimen, amethodof sequential load application shall be developed”bywhich each condition may, in turn, be tested to progressivelyhigher load levels. The final test may be taken to failure tosubstantiate the capability to accommodate internal loadredistribution, to provide data for any subsequent designmodification effort, and to provide data for use in any weightreduction programs. Failures at limit load shall includematerial yielding or deflection which degrade missionperforman~e and at ultimate load shall i~clude rupture orcollapse.

6.3.1.3 Test Levels and Duration

a. Static Koads. The loads, other than internalpressure in pressure vessels, shall be increaseduntil failure occurs or until the specified testloads are reached.

b. Temperature. Critical flight temperature-loadcombinations shall be used to determine theexpected worst case stress anticipated in flight.

c. Duration of Loadinq. Loads shall be applied asclosely as possible to actual flight loadingtimes, with a minimum dwell time sufficient torecord test data such as stress, strain,deformation, and temperature.

6.3.1.4 Supplementary Requirements. Pretest analysisshall be conducted to identify the locations of minimum designmargins and associated failure modes which correspond to theselected critical test load conditions. This analysis shall beused to locate instrumentation, to determine the sequence ofloading conditions, and to afford early indications of anomalousoccurrences during the test. This analysis shall also form thebasis for judging the adequacy of the test loads. Internalloads resulting from the limit test conditions shall envelop all

42



0

0

critical internal loads expected in flight; however, excessiveinternal loads peculiar to the test shall be avoided. In caseswhere a load or other environment has a relieving effect, theminimum, rather than the maximum, expected value shall be usedin defining limit test loading conditions. In some instances,where only a small number of flight vehicles have been includedin the program, the cost of a dedicated test article mayrepresent an unacceptably high percentage of the program cost.In such cases, the failure test would not be conducted, and itwould be necessary to subject flight hardware to test loadsprior to flight. In this event, special precautions shall betaken to ensure that the structure can still withstand itspredicted flight environment after it has been subjected to thetest loads. Such precautions shall include at least the specialdesign requirement that no permanent deformation detrimental tomission performance snail occur and the inspection requirementthat sufficient nondestructive testing be conducted after thetest to ensure the integrity of the structure. Alternatively,each flight vehicle shall be proof-tested; proof levels may beless than ultimate levels but shall exceed limit levels. Inthis case, the vehicle shall be designed to withstand the prooflevels without permanent deformation detrimental to missionperformance, and a thorough post-test inspection of each flightvehicle shall be conducted to ensure the integrity of thestructure.

6.4 COMPONENT QUALIFICATION TESTS

The space vehicle component qualification test baselineconsists of all the required tests specified in Table II. Thetest baSeline shall be tailored for each program, givingconsideration to both the required and optional tests; however,deviations from the baseline of required tests shall be approvedby the contracting officer. Sach component that is acceptancetested as a component shall undergo comparable qualificationtests as a component. Component qualification tests shallnormally be accomplished entirely at the component level.However, in certain circumstances, required componentqualification tests may be conducted partially or entirely atthe subsystem or space vehicle levels of assembly. Tests ofcomponents such as interconnect tubing, radio frequencycircuits, and wiring harnesaes are examples where at least someof the tests can usually be accomplished at higher levels ofassembly.

Where components fall into two or more categories of TableII, the required tests specified for each category shall beapplied. For example, a star sensor may be considered to fitboth ‘Electronic Equipment- and “Optical Equipment- categories.

43

8 A i


MIL-STD-1540B (USAF)10 OC’IXIBER1982

Table II. Component Qualification Tests

.*’3 & &sYs$ jg~ Ssg *. &

# @ * ‘# . ;$$~’ *= 2.$3 &

. .& @“$ G ges ● s% .$5 $$*“ * $’ 2* $$”$ Z* $=$$$S5$*

FUMCTIONAL8,4.1 ,(1) RR RRRR RRF!RR

THERMAL 6,a,2 9 RR RRRR RVhCUUM

ORRR

THERMAL 6,4.3 8CYCLING

Rrl o 0000–– ––

SINUSOIOAL6,4,4 5 00W6PA110N

0–0000 000

wrnu 6,4,5 4 RVMRAIION

~{3)R_RR8 ~RRR

ACOUSTIC 6,4.6 4 – R(31_R_— _—— ——

WROSHOCK 6.4.7 3 R o 00000 — 000”

ACCELEkA1lON6,4.8 7 0 R 00 o-— O––R

HUMIDITY 6,4.9 10 000000 0 0 000

PRESSURE 6,4.101, - – R – ~(2tnRRR __

LEAK 6,4.112,6,12 ,7(2) _ ~ — ~(2RRRoo _

EMC 6.4.1213 R 00–––– ––––

LIFE 8,4,1314000 000– 0 000

LEGENDR - REOU18EDO -,Oi’llONALTEST— - NOREQUIREMENT

Na18s11)FumU,mll"tismllti&uctmprtiBandtdlwwewi(onmen~,,e,(2)Wuirtionvo”wlw,rpr,s”,lzaoq,ipm,”,(3]EllherraWmtibratlOn~amSUCtestroq"l`e4tilhth80thOfwUml

In this example, a thermal cycling test would be conducted sinceit is required for electronic equ~gnnent, even though there is norequirement for thermal cycling optics. Similarly, an electricmotor-driven actuator fits both “Electrical Equipment” and“Moving Mechanical Assembly” categories. The former makesthermal cycling a required test, even though this test isoptional for the moving mechanical assembly category.

6.4.1 Functional Test, Component Qualification

6.4.1.1 Purpose. This test verifies that the electricaland mechanical performance of the component meets therequirements of the component specification.



I

6.4.1.2 Test Description. Electrical teats shall includeapplication of expected voltages, impedance, frequencies,pulses, and waveforms at the electrical interface of thecomponent, including all redundant circuits. Mechanical testsshall include application of torque, load, and motion asappropriate. These parameters shall be varied throughout theirspecification ranges and the sequences expected in flightoperation, and the component output shall be measured to verifythe component performance to specification KeqUIKeMentS.Functional performance shall alao include electrical continuity,stability, response time, alignment, pressure, leakage, or otherspecial functional tests related to a particular componentconfiguration.

6.4.1.3 Supplementary Requirement. A functional testshall be conducted prior to each of the environmental tests toassure that performance meets the requirements of the particularspecification. The same functional test shall be conductedafter each environmental test.

6.4.2 Thermal Vacuum Test, Component Qualification

6.4.2.1 Purpose. This test demonstrates the ability of

o

the component to perform in a thermal vacuum environment whichsimulates the design environment for the component (see 3.9).

6.4.2.2 Test Description. The component shall be mountedin a vacuum chamber on a thermally controlled heat eink or in amanner similar to its actual installation in the space vehicle.The component surface finishes shall be identical to those onthe flight models. A temperature sensor shall be attached tothe component baseplate for conduction-dominated internaldesigns or to a representative case location(s) for a componentcooled primarily by radiation. This sensor shall be used todetermine and control the test temperature. Temperaturestability has been achieved when the rate of change is no morethan 3 deg C/hour. The component heat transfer to the thermallycontrolled neat sink and the radiation heat transfer to theenvironment shall be controlled to the came proportions ascalculated for the flight environment. The components that arerequired to operate during ascent shall be operating andmonitored for arcing and corona during the initial reduction ofpressure to the specified lowest levels. The time for reductionof chamber pressure from ambient to 20 pascals (0.15 Torr) shallbe at least 10 minutes to allow sufficient time in the region ofcritical pressure. Components only operational during ascent:e:c:edturned off after the test pressure level has been

. Equipment that does not operate during launch, and

@

that may otherwise be subject to corona damage, shall have

45



electrical power applied after the test pressure level has beenreached. With the chamber at the test pressure level, RFequipment shall be monitored to assure that multipacting doesnot occur. A temperature cycle begins with the chamber atambient temperature. The temperature is reduced to thespecified low level and stabilized. All components whichoperate in orbit shall be turned off, then cold started after asoak period sufficient,to ensure the component internal

I temperature haS stabilized at the specified temperature, andthen functionally tested. With the component operating, thecom~nent temperature shall be increased to the uppertemperature level. After the component temperature hasstabilized at the specified level, the component shall be turnedoff, then hot started after electrical circuits have beendischarged, and then functionally tested. The temperature ofthe chamber shall then be reduced to ambient conditions. Thisconstitutes one complete temperature cycle.


a. Pressure. The pressure shall be reduced fromatmospheric to 0.0133 pascals (0.0001 Torr) orless.

b. Temperature. The component temperature shall beat the design high temperature during the hotcycle and at the design low temperature duringthe cold cycle.

c. Duration. A minimum of three temperature cyclesshall be used. Each cycle shall have a 12 houror longer dwell at the high and at the lowtemperature levels during which the unit isturned off until the temperature stabilizes andthen turned on.

6.4.2.4 SUP elementary Requirements. Functional testsshall be conducted at the high and low temperature levels duringthe first and last cycle and after return of the component toambient temperature. During the remainder of the test,electrical and electronic components, including all redundantcircuits and paths, shall be monitored for failures andintermittents to the maximum extent possible. Monitoring of theRF output for corona shall be conducted using spectrummonitoring instrumentation during chamber pressure reduction.The RF component shall be operated at maximum power and atdesign frequency. The force or torque design margin shall bemeasured on moving mechanical assemblies at the environmentalextremes. Compatibility of thrusters with their operational

46



.

0

0

fluids shall be verified at test temperature extremes duringthermal vacuum testing.

6.4.3 Thermal Cyclin9 Test, Component Qualification

6.4.3.1 Purpose. This test demonstrates the ability ofcomponents to operate over the design temperature range and tosurvive the thermal cycling screening test imposed upon thecomponent during acceptance testing.

6.4.3.2 Test Description. A thermal cycle begins with thecomponent at ambient temperature. With the component operating(power on) and while perceptive parameters are being monitored,the chamber temperature shall be reduced to the specified lowtemperature level as measured at a representative location onthe component such as the mounting point on the baseplate forconduction-dominated internal designs or at a representativelocation(s) on the case for radiation-controlled designs. Afterthe component temperature has stabilized at less than 3 degC/hour rate of change, the unit shall be turned off and thencold started. With the component operating, the chambertemperature shall be increased to the upper temperature level.After the component temperature has stabilized at the specifiedlevel, the component shall be turned off and then hot started.The temperature of the chamber shall then be reduced to ambientconditions. This constitutes one thermal cycle.


a. Pressure. Ambient pressure shall normally beused. When unsealed components are being tested,the chamber shall be flooded with dry air ornitrogen to preclude condensation on and withinthe component at low temperature. If convenient,this test may be performed in thermal vacuum andcombined with the test of 6.4.2, provided thatthe temperature limits, number of cycles, rate oftemperature change, and dwell times conform tothis test.

b. Temperature. The component temperatures shall beat the design high temeprature during the hotcycle and at the design low temperature duringthe cold cycle.

c. Duration. Three times the number of thermalcycles as used for acceptance testing but notless than 24 cycles total, of which the last 4shall be failure free. Each cycle shall have a

47



I

l-hour minimum dwell at the high and at the lowtemperature levels during which the unit shall beturned off until the temperature stabilizes andthen turned on. The dwell time at the high andlow levels shall be long enough to obtaininternal thermal equilibrium. The transitionsbetween low and high temperatures shall be at anaverage rate of ‘at least 1 deg C per minute.

6.4.3.4 Supplementary Requirements. Functional testsshall be conducted during the first and last thermal cycles athigh and low temperatures and after return of the component toambient. During the remainder of the test, electricalcomponents, including all redundant circuits, shall be cycledthrough various operational modes and perceptive parametersmonitored for failures and intermittents to the maximum extentpossible. Compatibility of valves and fluid or propulsionequipment with their operational fluids shall be verified attest temperature extremes during thermal cycling tests.

6.4.4 Sinusoidal Vibration Test, Component Qualification

6.4.4.1 Purpose. Sine wave vibration testing is used forone or more of the following purposes:

a. To demonstrate the ability of the component towithstand or, if appropriate, to operate at thedesign levels of the sinusoidal or decayingsinusoidal type dynamic vibration environmentspecified for the component.

b. To determine any resonant conditions which couldresult in failure in flight or in subsequentvibration tests.

c. ‘lbevaluate fixtures or for diagnostic purposes.

6.4.4.2 Test Description. The component shall be mountedto a fixture through the normal mounting points of thecomponent. The component shall be tested in each of threemutually perpendicular axes. Significant resonant frequenciesshall be noted and recorded. The induced cross axisaccelerations at the attach points should be limited to themaximum test levels specified for the cross axes.

6.4.4.3 Test Levels and Duration. Tests conducted todetermine resonant conditions or to evaluate fixtures (purposesOb. or “cm above) shall be conducted Ue.lngtest levels anddurations sufficient to provide diagnostic capability.Sinusoidal excitation may be applied as a dwell at discretefrequencies or as a frequency sweep with the frequency varyingat a logarithmic rate. The sweep rate for diagnostic tests

48

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o,.I

I

shall be slow enough to allow identification of significantresonances. Tests conducted to demonstrate the degree ofruggedness (purpose “a” above) shall use two (2) minutes peroctave unless the sweep rates and dwell times can be based onthe persistence of the environment in service use. Thevibration levels shall be sufficient to cover the severity ofthe maximum design levels.

6.4.4.4 Supplementary Requirements. Sinusoidal vibrationtests to demonstrate the degree of ruggedness (purpose “a”above) shall be considered a required test where significantsinusoidal vibration is expected in service usage. A functionaltest shall be conducted before the sinusoidal vibration test andafter ita completion. Electrical components shall be poweredduring the test and perceptive parameters monitored for failuresor intermittent. When monitoring during the test ia notpractical, a limited functional test shall be performed afterthe vibration test for each axis is completed. If the componentis to oe mounted on dynamic isolators in the space vehicle, thecomponent shall be mounted on these isolators during thequalification test.

6.4.5 Random Vibration Test, ComPonent Qualification

6.4.5.1 Purpose. ThiS test demonstrates the ability ofthe component to withstand the design level random vibrationenvironment.

6.4.5.2 Test Description. The component shall be mountedto a rigid fixture through the normal mounting points of thecompnent. The component shall be tested in each of threemutually perpendicular axes. Propulsion system valves shall bepressurized to operating pressure for this test and monitoredfor internal pressure decay if pressurized during ascent.

6.4.5.3 Test Levels and Duration. The random vibrationshall be the design level for the component. The minimumoverall test level for components weighing 22.7 kg (50 lb) orless shall be 12 grins. The minimum teSt level fOr componentsweighing more than 22.7 kg (50 lb) shall be evaluated on anindividual basis. Tne test duration in each of the threeorthogonal axes shall be three times the expected flightexposure time to the maximum predicted environment or threetimes the component random vibration acceptance test time ifthat is greater, but not leas than 3 minutes per axis. Whereinsufficient time is available at the full test level to testall redundant circuits, all functions, and all modes, extendedtesting at a level 6 dB lower shall be conducted as necessary tocomplete functional testing.

49



6.4.5.4 SUpplementary Requirements. A functional testshall be conducted before and after the completion of the randomvibration test. During the test, electrical and electroniccomponents, including all redundant circuits, shall beelectrically energized and functionally sequenced throughvarious operational modes to the maximum extent possible.Several perceptive parameters shall be monitored for failures orintermittents during the test. If the component is to bemounted on dynamic isolators in the space vehicle, the componentshall be mounted on these isolators during the qualificationtest and vibration test levels controlled at the input to theisolators.

6.4.6 Acoustic Test, Component Qualification

6.4.6.1 Purpose. This test demonstrates the ability ofthe component to withstand the design level acousticenvironment. Acoustic tests shall be conducted only oncomponents with large surfaces which are likely to besusceptible to acoustic noise excitations.

6.4.6.2 Test Description. The component shall beinstalled in a reverberant acoustic cell capable of generatingdesired sound pressure levels. A uniform sound energy densitythroughout the chamber is desired. The configuration of thecomponent, such as deployed or stowed, shall be as it is duringsubjection to the flight dynamic environment. The preferred o

method of testing shall be with the component mounted onflight-type support structure and with ground handling equipmentremoved. Electrical components shall be operating.

6.4.6.3 Test Levels and Duration. The sound pressurelevel shall be at least the design level, but not less than 144dB overall. The duration shall be three times the expectedflight exposure time to the maximum predicted environment orthree times the acoustic acceptance test duration, whichever isgreater, but not less than three minutes. Where insufficienttime is available at the full test level to test all redundantcircuits, all functions, and all modes, extended testing at alevel 6 dB lower shall be conducted as necessary to completefunctional testing.

6.4.6.4 Supp lementary Requirements. A functional testshall be conducted before and following the acoustic test.During the test, electrical and electronic components, includingall redundant circuits, shall be electrically energized andfunctionally sequenced through various operational modes to themaximum extent possible. Several perceptive parameters shall bemonitored for failures or intermittents during the test.

50


MIL-sTD-1540B (USAF)10 OCTOBER 1982

Components characterized by large ratios of surface area tovolume, such as large antennas and solar arrays, cannot beteated in a manner which suitably simulates imposition of theservice dynamic.environment by employing mechanical vibration.For such component configurations acoustic testing shall berequired. When acoustic component testing is required, randomvibration component testing shall not be required.

6.4.7 Pyro Shock Test, Cornponent Qualification

6.4.7.1 Purpose. This test demonstrates the capability ofthe component to withstand the design level pyrotechnic shockenvironment.

6.4.7.2 Test Description. The component shall be mountedto a fixture through the normal mounting points of thecomponent. The selected test method shall be capable of meetingthe required shock spectrum with a transient which has aduration comparable to the duration of the expected inflightshock . Numerous test techniques are currently in use for the

as

performance of pyrotechnic s~ock testing. Me_ihods such asshaped pulses, complex decaying sinusoids, impact devlces~ andpyrotechnically excited fixtures have been successfullyemployed. Electrodynamics shakers with control electronicscapable of synthesizing the shock transient with a prescribedspectrum have been developed. A mounting of the equipment onactual or dynamically similar structure provides a morerealistic test than does a mounting on a rigid structure sucha shaker armature or slip table. A rigid structure tends tooverly correlate the input shock motions at the equipmentattachment points and thereby can lead to higher responseswithin the equipment than intended. Sufficient developmenttesting shall be conducted to validate the proposed test methodbefore testing qualification hardware. These test techniques aswell as others that may be devised are acceptable, provided thefollowing conditions are met:

a. A transient with the prescribed shock spectrumcan be generated within specified tolerances.

b. The applied shock transient provides asimultaneous application of the frequencycomponents as opposed to a serial application.Toward this end, it shall be a goal for theduration of the shock transient to approximatethe duration of the actual shock event. Ingeneral, the duration of the shock employed forthe shock spectrum analysis shall not exceed 20

51


MIL-STD-1540B (USAF)10 OC213BER1982

I

milliseconds unless a longer duration is required ●I to simulate the actual shock event.

I 6.4.7.3 Test Levels and Exposure. The shock spectrum ineach direction along each of the three orthogonal axes shall be

~

at least the design level for that direction. A sufficientnumber of shocks shall be imposed to meet the amplitude criteriain both directions on each of the three orthogonal axes at leastthree times. However, if a suitable test environment can begenerated to satisfy the amplitude requirement in all six axialdirections by a single application, this test environment shallbe imposed three times. If, on the other hand, an imposed shockmeets the amplitude requirement in only one direction of asingle axis, the shock test shall be conducted a total of 18times in order to get three valid test amplitudes in bothdirections of each axis.

6.4.7.4 SUP elementary Requirements. A visual inspectionshall be made before and after the test. The visual inspectionshall not entail the removal of component covers nor anydisassembly. Electrical and electronic components, includingredundant circuits,. shall be energized and monitored to themaximum extent possible. A functional test shall be performedbefore and after all shock tests, and several perceptiveparameters monitored during the shocks to evaluate performanceand to detect any failures. Relays shall not transfer and shallnot chatter in excess of specification limits. If the component ●is to be mounted on dynamic isolators in the space vehicle, thecomponent shall be mounted on these isolators during thequalification test.

6.4.8 Acceleration Test, ComPonent Qualification

6.4.8.1 Purpose. This test demonstrates the capability ofthe component to withstand or, if appropriate, to operate in thedesign level acceleration environment.

6.4.8.2 Test Description. The component shall be mountedto a test fixture through the normal mounting points of thecomponent. The component shall be tested in each of threemutually perpendicular axes. The specified accelerations applyto the geometric center of the test item. If a centrifuge isused, the arm (measured to the geometric center of the testitem) shall be at least five times the dimension of the testitem measured along the arm. Inertial components such as gyrosand platforms may require counter rotating fixtures on thecentrifuge arm.


52



0 .a. Acceleration L@V@l. The test acceleration level

shall be at least the design levels (see 3.12) ineach direction for each of the three orthogonalaxea. For the axis in the direction of thelaunch acceleration, the test level shall not beless than 20 g for tests qualifying componentsfor the launch and injection environments.

b. Duration. Five minutes each axis in eachdirection.

6.4.8.4 Supplementary Requirements. A functional testshall be conducted before the acceleration test and aftercompletion of the test. Electrical components, if operatedduring flight, shall be powered during the test and perceptiveparameters monitored for failures or intermittent. If thecomponent is to be mounted on dynamic isolators in the spacevehicle, the component shall be mounted on these isolatorsduring the qualification test.

6.4.9 Humidity Test, Component Qualification

6.4.9.1 Purpose. This test demonstrates that thecomponent is capable of surviving without excessive degradationthe design value of humidity which might be imposed upon thecomponent during fabrication, test, shipment, storage, andpreparation for launch.

6.4.9.2 Test Description and Levels. The component shallbe installed in the chamber.

a. Pretest Conditions. Chamber temperature shall beat room ambient conditions with uncontrolledhumidity.

b. ~. The temperature shall be increased to+35 deg C over a l-hour period; then the humidityshall be increased to not less than 9S percentover a l-hour period.with the temperaturemaintained at +35 deg C. These conditions shallbe held for 2 hours. The temperature shall bereduced to 2 deg C over a 2 hour period with therelative humidity stabilized at not less than 95percent. These conditions shall be held for 2hours.

c. SYskQ. The above cycle shall be repeatedexcept that the temperature shall be increasedfrom +2 deg C to +35 deg C over a 2 hour period

53



(moisture is not added to the chamber until +35deg C is reached).

d. ~. The chamber temperature shall beincreased to +35 deg C over a 2 hour periodwithout adding any moisture to the chamber. Thetest component shall then be dried with air atroom temperature and 50 percent maximum relativehumidity by blowing air through the chamber for 6hours. The volume of air used per minute shallbe equal to one to three,times the test chambervolume. A suitable container may be used inplace of the test chamber for drying the testcomponent.

e. SZEQ. The component shall be placed in thetest chamber and the temperature increased to +35deg C and the relative humidity increased to 90percent over a 1 hour period and then these endconditions shall be maintained for at least 1hour. The temperature shall be reduced to +2 degC over a 1 hour period with the relative humiditystabilized at 90 percent and these conditionsmaintained for at least 1 hour. A drying cycleshould follow (see Cycle 3).

6.4.9.3 SUP elementary Requirements. The component shallbe functionally tested prior to the test and at the end of Cycle3 (within 2 hours after the drying) and visually inspected fordeterioration or damage. The component shall be functionallytested during the Cycle 4 periods of stability; after the l-hourperiod to reach +35 deg C and 90 percent relative humidity, andagain after the l-hour period to reach the +2 deg C and 90percent relative humidity condition. The component shall bevisually inspected for cieterioration or damage after removalfrom the chamber.

6.4.10 Pressure Test, Component Qualification

6.4.10.1 Purpose. This test demonstrates that the designand fabrication of such items as pressure vessels, pressurelines, fittings, and valves provides an adequate margin suchthat structural failure or excessive deformation does not occurat the maximum expected operating pressure.

6.4.10.2 Test Description

●

a. Proof Pressure. For such items as pressurevessels, pressure lines, and fittings, the

54


IMIL-STD-1540B (USAF)10 OCTOBER 1982

0 temperature of the component shall be consistentwith the critical use temperature and subjectedto a minimum of one cycle of proof pressure. Aproof pressure cycle shall consist of raising theinternal pressure (hydrostatically orpneumatically, as applicable) to the Proofpressure, maintaining it for 5 minutes and thendecreasing the pressure to zero. Evidence ofpermanent set or distortion that exceeds 0.2percent or failure of any kind shall indicatefailure to pass the test.

b. Proof Pressure for Valves. With the valve in theopen and closed positions (if applicable), theproof pressure shall be applied for a minimum ofthree cycles to the inlet port for 5 minutes(hydrostatically or pneumatically, asapplicable) . Following the 5-minutepressurization period, the inlet pressure shallbe reduced to ambient conditions. The exteriorof the unit shall be visually examined. Evidenceof deformation thst exceeds 0.2 percent or anyfailure shall indicate failure to pass the test.The test may be conducted at room ambienttemperature.

c. Burst Pressure (see 3.4). For such items aspressure vessels, pressure lines, and fittings,the temperature of the component shall beconsistent with the critical use temperature, andthe component shall be pressurized(hydrostatically or pneumatically, as applicableand safe) to design burst pr&sure or greater.The internal pressure shall be applied at auniform rate such that stresses are not imposeddue to shock loading.

d. Burst Pressure for Valves. With the valve in theopen or closed posltlon, as applicable, thedesign burst pcessure shall be applied to theinlet port for 5 minutes (hydrostatically orpneumatically, as applicable). Following the5-minute pcessucization period, the inletpressure shall be reduced to ambient conditions.The exterior of the unit shall be visuallyexamined for indications of deformation orfailure. The test may be conducted at roomambient temperature.

55


MIL-STD-1S40B (USAF)10 OC’lUBER 1982

6.4.10.3 Test Levels

I

a. Temperature. AS specified in the testdescription. As an alternative, tests may beconducted at ambient room temperatures if thetest pressures are suitably adjusted to accountfor temperature effects on strength and fracturetoughness.

b. Proof Pressure. Unless otherwise specified, theproof pressure equals 1.5 times the maximumoperating pressure.

c. Burst Pressure. Unless otherwise specified, theburst pressure equals two times the maximumoperating pressure.

6.4.10.4 Supplementary Requirements. The component shallwithstand proof pressure without leakage or detrimentaldeformation. Applicable safety standards shall be followed inconducting all tests.

6.4.11 Leakage Test, Component Qualification

6.4.11.1 Purpose. This test demonstrates the capabilityof pressurized components to meet the design leakage rateconstraints specified in the component specifications.

●6.4.11.2 Test Description and Alternatives. Component

leak checks shall be made prior to initiation of, and followingthe completion of, component qualification thermal and vibrationtests. Proof pressure tests per 6.4..10shall be successfullycompleted before conducting leakage tests. The test methodemployed shall have sensitivity and accuracy consistent with thespecified maximum allowable leak rate. One of the followingrecommended methods shall be used:

a. Method I (9ross leak test). The component shallbe completely immersed in a liquid so that theUPPeKmOSt Part of the test item is 5 ~ 2.s cm (2~ 1 inches) below the surface of the liquid. Thecritical side or side of interest of thecomponent shall be in a horizontal plane facingUP . The liquid, pressurizing gas, and the testitem shall be 23 ~ 10 deg C (73 ~ 18 deg F). Thegas used for pressurizing shall be clean and drywith a dewpoint of at least -32 deg C (-25 degF). MY observed leakage during immersion asevidenced by a continuous stream of bubbles

I56

I



e

i-

b.

c.

d.

e.

emanating from the component indicates a failureof seals.

Method II (fine leak test)_. The component shallbe purged with nitrogen and then charged withhelium to the required pressure (as specified inthe component detail specification) before beingsealed. The component shall then be placed in asuitable vacuum chamber and tested for heliumleakage with a helium leak detector. The leakagerate shall be used to determine seal integrityand shall not exceed the amount specified in thedetailed component specification. This method isapplicable to tape recorders and similarcomponents.

Method III (for battery cases or pressurized

-~roqen or other appropriate gas to theThe component shall be pressurized

specified valu~. The pressuie shall be-monitoredby a gage (or pressure transducer) for therequired time. The drop in pressure shall notexceed the permitted amount as specified underthe component specification.

Method IV (for hermetically sealed alkalinestorage batteries). The battery shall be cleanedwith alcohol while in the discharged state. Asuitable indicator (e.g., dilute solution ofphenolphthalein or other suitable color changeindicator) shall be applied to all seams,terminals, and pinch tubes subject to leakage ofelectrolyte. A change in the color of theindicator shall be an indication of a leak.After testing, the test solution shall be removed(e.9., with distilled water).

Method V (for components of pressurized fluid

~“ The components shall be pressurized totheir maximum working pressure in each of thefunctional modes. ~a~age shall be detectedusing an appropriate method (6.2.6.4).Propulsion system tanks and thrusters shall alsobe evacuated to the internal pressure normallyused for propellant loading and the internalpressure monitored for indications of leaking.

6.4.11.3 Test Levels and Duration, .The leak tests shallbe performed with the component pressurized at the maximum

57


MIL-STD-1540B (uSAP)10 OC’IOBER1982

operating pressure and then at the minimum operating pressure if ●the seals are dependent upon pressure for proper sealing. Thetest duration shall be sufficient to detect any significantleakage. The test levels and duration for the typical methodsof 6.4.11.2 are:

a. Method I. The duration of immersion shall be 60minutes at each pressure.

b. Method II. The external test pressure shall be0.133 pascals (0.001 Torr) or less and theduration of the test shall be 4 hours (forequipment that is operational in orbit for morethan one day).

c. Method III. The test pressure is usually lessthan 343 kilopascals (50 psi). The pressure dropshall not exceed the specified amount (typicallyabout 6.9 kilopascals (1 psi) in a 6-hour periodat room temperature).

d. Method IV. The test results are VlSdbl(3 withinseconds.

e. Method V. The duration of the evacuatedpropulsion system component leak test shall notexceed the time that this condition is normallyexperienced during propellant loading.

6.4.11.4 Supplementary Requirements. Component leak testsare considered adjunctive to the component qualificationenvironmental tests in that their results are part of thesuccess criteria for these tests.

6.4.12 EMC Test, Component Qualification

6.4.12.1 Purpose. This test demonstrates that theelectromagnetic interference characteristics (emission andsusceptibility) of the component under normal operatingconditions does not result in malfunction of the component andthat the component does not emit, radiate, or conductinterference which results in malfunction of other systemcomponents.

6.4.12.2 Test Description. The test shall be conducted inaccordance with the requirements of MIL-STD-1541. An evaluationshall be made of each component to determine which tests shallbe performed as the baseline requirements.

o58


Io


6.4.13 Life Test, ComPonent Qualification

6.4.13.1 PU~POS:. This test demonstrates the reliabilityof the component and Increases confidence that components whichmay have a wearout, drift, or fatigue-type failure mode have thecapability to withstand the maximum duration or cycles ofoperation to which they are expected to operate during repeatedground testing and in flight without degradation of theirfunction outside of allowable limits.

6.4.13.2 Test Description. One or more components shallbe set up to operate in conditions that simulate the flightconditions to which they would be subjected. Theseenvironmental conditions shall be selected for consistency withend use requirements and the significant life characteristics ofthe particular component. Typical environments are ambient,thermal, thermal vacuum, and various combination of these. Thetest sample shall be selected at random from production units orshall be a qualification unit. The test shall be designed todemonstrate the ability of the component to withstand themaximum operating time and the maximum number of operationalcycles predicted during its service life with a suitablemargin. For components having a relatively low percentage dutycycle, it shall be acceptable to compress the operational dutycycle into a tolerable total teat duration. For componentswhich operate continuously in orbit, or at very high percentageduty cycles, accelerated test techniques may be employed if suchan approach can be shown to be valid.

6.4.13.3

a.

b.

c.

Test Levels and Duration

Pressure. Ambient pressure shall be used exceptfor unsealed units where degradation due to avacuum environment may be anticipated. In thosecases, a pressure of 0.0133 pascals (0.0001 Torr)or less shall be used.

Environmental LeVelS. The maximum predictedenvironmental levels shall be used. Foraccelerated life teata, environmental levels maYbe selected that are more severe than flightlevels, provided the higher stresses can becorrelated with life at the predicted usestresses and do not introduce additional failuremechanisms.

Duration. The total operating time or number ofoperational cycles for a component life testshall be twice that predicted during the service

59



life, including ground testing, in order todemonstrate an adequate margin.

d. Functional Duty @cle. Complete functional testsshall be conducted before the test begins, aftereach 168 hours of operation and during the last 2hours of the test. ti abbreviated functionaltest shall be conducted periodically to aacertainthat the component is functioning withinspecification limits.

6.4.13.4 Supplementary Requirements. For statistical typelife tests, the duration is dependent upon the number ofsamples, confidence, and reliability to be demonstrated.

6.5 SUBASSEMBLY LEVEL QUALIFICATION TESTS

Subassembly level qualification tests shall be conducted onthose subassemblies that are subjected to environmentalacceptance tests at the subassembly level. For othersubassemblies, qualification tests are to be considered asoptional unless specified otherwise in the contract. Functionalor environmental qualification tests may be conducted at thesubassembly level to detect material and workmanship defects, orto measure critical parameters, that cannot be accomplishedsatisfactorily at higher levels of assembly. When subassemblylevel qualification tests are planned, the subassemblies may be ●tested to similar requirements as components, or if morestringent requirements are used for acceptance test stressscreening then the more stringent levels shall be the basis forthe qualification tests. In general, all parts shall bequalified to maximum and minimum environmental levels well inexcess of the levels predicted for their specific application inthe space vehicle.

60


IMIL-STD-1540B (USAF)10 OCTOBER 1982 ,

●

,

0

0

SECTION 1

ACCEPTANCE TESTS

7.1 SPACE VEHICLE ACCEPTANCE TESTS

The space vehicle acceptance test baseline consists of allthe required tests specified in Table III. The teSt baselineshall be tailored for each program, giving consideration to boththe required and optional tests; however.,deviations from thebaseline requirements for the required tests shall be approvedby the contracting officer.

Table III. Space Vehicle Acceptance Tests

Reference Suggested Required (R)Test Paragraph Sequence OR

Optional (0)

Functional 7.1.1 ~(1) R

Snc 7.1.2 2 0

Acoustic 7.1.3 5 R(2)

Vibration 7.1.4 5 0

Pyro Shock 7.1.5 4 0

Pressure 7.1.6 3, 6 R

Thermal Vacuum 7.1.7 8R(3)

Thermal Cycling 7.1.8 7 0

Storage Tests 7.1.9 0

Special Tests 7.1 0

Notes: (1) Electrical functional tests shall be conductedprior to and following each environmental test.

(2) Conduct vibration in place of acoustic test forvehicles of compact shape and with weight lessthan 180 kilograms.

(3) Requirements are modified if Thermal Cycling test7.1.8 is conducted.

61

I


MIL-STD-1540B (USAF).10 OC’lX)BER1982

Additional special tests normally conducted by spacevehicle programs include alignments, instrumentationcalibrations, and measurements of mass properties, antennapatterns, and magnetic field. Since performance and accuracyrequirements are generally program peculiar, and test methodsare typically contractor peculiar, these tests are not includedin this standard.

If the space vehicle is controlled by on-board dataprocessing, the flight version of the computer software shall beresident in the space vehicle computer for these tests. Theverification of the operational requirements shall bedemonstrated in these tests to the extent practicable.

7.1.1 Functional Test, Space Vehicle Acceptance

7.1.1.1 Purpose. This test verifies that the electricaland mechanical performance of the space vehicle meets theperformance requirements of the specifications and detects anyanomalous condition.

7.1.1.2 Mechanical Functional Test. Same as themechanical functional test for space vehicle qualification(6.2.1.2), except tests are only necessary at nominaloperational conditions.

7.1.1.3 Electrical Functional Test. Same as theelectrical functional test for space vehicle qualification(6.2.1.3), except tests are only necessary at nominaloperational conditions.

The final ambient functional test conducted prior toshipment of the space vehicle to the launch base provides thedata to be used as success criteria during launch base testing.For this reason, the functional test should be designed so thatit can be duplicated, as nearly as possible, at the launchbase. The results of all factory functional tests and of thoseconducted at the launch base shall be used for trend analysis.

7.1.1.4 Supplementary Requirements. Same as 6.2.1.4.

7.1.2 EMC Test, Space Vehicle Acceptance. Limited EMCacceptance testing shall be accomplished on space vehicles tocheck on marginal EMC compliance indicated during space vehicleEMC qualification testing and to verify that major changes havenot occurred on successive production equipment. The limitedtests shall include measurements of power bus ripple, peaktransients, and monitoring of selected critical circuitparameters.

62



I

I

7.1.3 Acoustic Test, Space Vehicle Acceptance

7.1.3.1 Purpose. This test simulates the acoustic andvibration environment imposed on a space vehicle in flight inorder to detect material and workmanship defects that might notbe detected in a static test condition.

7.1.3.2 Test Description. Same as 6.2.3.2.

7.1.3.3 Test Levels and Duration. The acoustic spectrumshall represent the maximum predicted flight environment asdefined in 3.20. The overall sound pressure level foracceptance testing.shall not be less than 138 dB. The exposuretime at full acceptance test level shall equal or exceed themaximum expected flight exposure time, but the test time shallnot be less than 1 minute. Operating time should be dividedapproximately equally between redundant circuits. Whereinsufficient time is available at the full test level to testall redundant circuits, all functions, and all modes, extendedtesting at a level 6 dB lower shall be conducted as necessary tocomplete functional testing.


7.1.4 Vibration Test, Space Vehicle Acceptance

7.1.4.1 Purpose. This test simulates the dynamicvibration environment imposed on a vehicle in flight in order todetect material and workmanship defects. This test is only usedfor small vehicles with weights under 150 kg and compact shapes.


7.1.4.3 Test Levels and Duration. Random vibration testlevels shall be at the maximum predicted environmental levels asdefined in 3.23. Test duration shall be equal to or exceed theexpected flight exposure time to the maximum predictedenvironment but not less than 1 minute in each of three axes.Different axes may have different levels applied. Sinusoidaltest levels shall be at the maximum predicted environmentallevels as defined in 3.24, and test duration shall be equal tothe expected flight exposure to the makimum predictedenvironment. Where insuffi~ient time is available at the fulltest level to test all redundant circuits, all functions, andall modes, extended testing at a level 6 dB lower shall beconducted as necessary to complete functional testing.

7.1.4.4 SUPP lementary Requirements. Same as 6.2.4.4.

63


MIL-STD-1540B (USAF’)10 OCTOBER 1982

7.1.5 Pyro Shock Test, Space Vehicle Acceptance

7.1.5.1 Purpose. This test simulates the dynamic shockenvironment imposed on a space vehicle in flight in order todetect material and workmanship defects.

I7.1.5.2 Test Description. Same as 6.2.5.2.

7.1.5.3 Test Levels and Duration. Pyrotechnic shockacceptance testing of space vehicles shall be required in thoseinstances where the shock-producing mechanism can be readilyrefurbished for flight, as is often the case for explosive nuts,bolts, pinpullers, and clamps. One firing of those pyrotechnicdevices causing significant shocks to critical and shocksensitive components shall be conducted. Firing of both primaryand redundant pyres is required in the same relationship as theywill be used in flight. However, where the pyrotechnicmechanism explosively severs structure by detonation ofdetonating fuse or shaped charge, such testing shall not beincluded or required. To aid in fault detection, the pyro shocktest shall be conducted with subsystems operating and monitoredto the maximum extent practical.

7.1.6 Pressure Test, Space Vehicle Acceptance

7.1.6.1 Purpose. This test demonstrates the capability offluid subsystems to meet the flow, pressure, and leakaqe ●requirements

7.1.6.2

7.1.6.3be performed

specified in the space-vehicle- specification.

Test Description. Same as 6.2.6.2.

Test Levels and Duration. The leak checks shallby pressurizing the subsystem to maximum operating

pressure and holding at this pressure for a period commensuratewith the leakage method being employed.

7.1.6.4 5UPP lementary Requirements. Same as 6.2.6.4.

7.1.7 Thermal Vacuum Test, Space Vehicle Acceptance

7.1.7.1 Purpose. This test detects material, process, andworkmanship defects that would respond to thermal vacuum andthermal stress conditions and verifies thermal control.


7.1.7.3 Test Levels and Duration. Temperatures in variousequipment areas shall ‘be controlled by the external testenvironment and internal heating resulting from equi~ent

64



0operation so that the hot (or cold) temperature on at least onecomponent in each equipment area equals the maximum (or minimum)predicted temperature as defined in 3.25. The temperatureextremes shall be established by a survey of predictedtemperatures in various equipment areas and may have to beadjusted to performance of the most sensitive components in aparticular area. The pressure shall be maintained at 0.0133pascals (0.0001 Torr) or less. Duration shall be sufficient totest all orbital operational conditions and all equi~entfunctional modes including redundancy. Operating time should be .divided approximately equally between redundant circuits. Ifthe thermal cycling test (7.1.8) is not conducted, the thermalvacuum acceptance test shall include at least four completehot-cold cycles at the maximum predicted orbital rate oftemperature change and have at least an 8-hour soak at eachtemperature extreme of each cycle.

During one temperature cycle, thermal equilibrium shall beachieved at both hot and cold extremes to allow verification ofperformance of the thermostats, louvers, heat pipes, electricheaters, and the control authority of active thermal systems.Thermal equilibrium has been achieved when equipment temperaturechange is not more than 3 deg C/hour.

7.1.7.4 SUPPlementary Requirements. It may be necessaryto force temperature extremes at certain locations by alteringthermal boundary conditions locally or by altering theoperational sequence to provide additional heating or cooling.Any special conditioning within the space vehicle shallgenerally be avoided. External baffling, shadowing, or heatingshall be utilized to the extent possible.

7.1.8 Thermal Cycling Test, Space Vehicle Acceptance

7.1.8.1 Pur+“

This test detects material, process, andworkmanship de ects by subjecting the space vehicle to a thermalcycling environment.


7.1.8.3 Test Levels and Duration. The space vehicletemperature range from hot to cold shall be the maximum possiblewithin the constraints of the components acceptancetemperatures. The minimum space vehicle temperature range shallbe 50 deg C. Auxiliary heating and cooling may be used toprotect selected temperature sensitive components. The averagerate of change of temperature from one extreme to the othershall be as rapid as possible. Operating time should be dividedapproximately equally between redundant circuits. The minimum

65



I

I

number of thermal cycles shall normally be 40.

7.1.8.4, Supplementary Requirements. If this test isimplemented, only one thermal cycle is required in the thermalvacuum acceptance test s~ecified in 7.1.7. Consideration shouldbe given to conducting this test where considerable disassemblyfor rework of components has occurred or if maximum confidencein the system is required.

7.1.9 Storage Tests, Space Vehicle Acceptance. The spacevehicle is sometimes stored at the contractor’s facility forseveral months until needed for replenishment of an on-orbitoperational system. Careful planning of environmentalconditions and any testing during storage is necessary.Systematic checks of the space system health shall be made atperiodic intervals while in storage. Age-sensitive parts shallbe accounted for and reported on a schedule basis. Componentshaving rotating elements may require periodic operation. If thespace vehicle is removed from storage and design modificationsmade, consideration shall be given to conducting all of therequired acceptance tests specified in Table III. Upon thevehiclets removal from storage for shipment to the launch site,a complete ambient functional acceptance test shall be conductedprior to shipment.

7.2 SUBSYSTEM ACCEPTANCE TESTS●

Subsystem level acceptance tests are considered optional.These tests are often cost-effective because failures detectedat this level usually are less costly to correct than are thosedetected at the space vehicle level, Acceptance tests should beconducted at the subsystem level where this level provides amore perceptive test than would be possible at the space vehiclelevel. The desirability of conducting these subsystem

acceptance tests should be evaluated considering such factors asthe relative accessibility of the subsystem and its componentand the retest time at the vehicle level. The possiblenecessity to repeat other portions of vehicle level testsbecause of their invalidation due to a subsystem failurecorrective action should also be considered. When subsystemlevel tests are performed, the test requirements shall be basedon the space vehicle level test requirements.

7.3 COMFQNENT ACCEPTANCE TESTS

The space vehicle component acceptance test baselineconsists of all the required tests specified in Table IV. Thetest baseline shall be tailored for each program, givingconsideration to both the required and optional tests; however,

@

66


MIL-STD-1540B (USAF)10 OCTQBER 1982

●

I,-

1

I

0

TABLE IV. Component Acceptance Tests

THWIAL 7.3,2 7 R(2I ~ R ORRR OR FIRVAWUM

THERMAL 1.3.3 6 R 0 0 0 oo o---–rfcuffi

KuMM 1,3,4 4 R ~4)R-o$IRo RRRVlmlAImx

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deviations from the baseline of required tests shall be approvedby the contracting officer. Component acceptance tests si;llnormally be accomplished entirely at the compOnent level.However, in certain circumstances, the required componentacceptance tests may be conducted partially or entirely at thesubsystem or space vehicle levels of assembly. Acceptance teStSof components such as interconnect tubing, radio frequencYcircuits, and wiring harnesses are examples where at least someof the tests can usually be accomplianed at higher levels ofassembly.

Where components fall into two or more categories ofTable IV, the required tests specified for each category shallbe applied. For example, a star sensor may be considered t? fitbth “Electronic Equipment” and “Optical Equipment” Categories.In this example, a thermal cycling test would be conducted since

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MIL-STD-1540B (USAF)10 OC!lKIBER1982

it is required for electronic equipment, even though there is norequirement for thermal cycling optics. Similarly, an electricmotor-driven actuator fits both “Electrical Equipment” and“Moving Mechanical Assembly” categories. The former makesthermal cycling a required test, even though this test isoptional for the moving mechanical assembly category.

7.3.1 Functional Performance Test, Component Acceptance

7.3.1.1 Purpose. This test verifies that the electricaland mechanical performance of the component meets the specifiedoperational requirements of the component.



7.3.2 Thermal Vacuum Test, Component Acceptance

7.3.2.1 Purpose. This test detects material andworkmanship defects ~rior to installation into a space vehicle,by subjecting the unit to a thermal vacuum environment.


7.3.2.3 Test Levels/Duration

a. Pressure. The pressure shall be reduced fromatmospheric to 0.0133 pascals (0.0001 Torr) orless.

b. Temperature. The high temperature shall be themaximum predicted but not,less than +61 deg C,and the low temperature shall be the minimumpredicted temperature but not higher than -24 degC, except where the temperature extreme wouldresult in physical deterioration of the materialsin the component such as in tape recorders andbatteries.

c. Duration. A minimum of one temperature cycleshall be used. The cycle shall have a 12-hourdwell at the high and at the low temperaturelevels during which the unit is turned off untilthe temperature has stabilized and then turned on.


7.3.3 Thermal Cycling Test, Component Acceptance

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MIL-STD-1540B (uSAP)~ 10 OCTOBER 1982

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7.3.3.1 Purpose. This test detects material andworkmanship defects prior to installation of the component intoa space vehicle, by subjecting the component to thermal cycling.



a. Pressure. Ambient pressure shall normally beused. When unsealed components are being tested,the chamber shall be flooded with dry air ornitrogen to preclude condensation on and withinthe component at low temperature. If convenient,this test may be performed in thermal vacuum andcombined with the teat of 7.3.2, provided thatthe temperature limits, number of cycles, rate oftemperature change, and dwell times conform tothis test.

b. Tempe rature. The high temperature Shall be themaximum predicted but not less than +61 deg C,and the low temperature shall be the minimumpredicted but not higher than -24 deg C, exceptwhere the temperature extreme would result inphysical deterioration of the materials in thecomponent such as in tape recorders and batteries.

c. Duration. The minimum number of temperaturecycles shall be eight, of which the last fourshall be failure free. Each cycle shall have al-hour minimum dwell at the high and at the lowtemperature levels du,ringwhich the unit shall beturned off until the temperature stabilizes andthen turned on. The dwell time at the high andlow levels shall be long enough to obtaininternal thermal equilibrium. The transitionsbetween low and high temperatures shall be at anaverage rate of at least 1 deg C per minute.


7.3.4 Random Vibration Test, Component Acceptance

7.3.4.1 Purpose. This test detects material andworkmanship de~ects prior to installation of the component intoa flight system, by subjecting the unit to a dynamic vibrationenvironment.

7.3.4.2 Teat Description. Same as 6.4.5.2.

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7.3.4.3 Test Levels and Duration. The vibration testspectrum shall be equal to or greater than a smooth spectralrepresentation for the maximum predicted environment as definedin 3.23: ~he minimum overall test level for components weighing22.7 kg (50 lb) or less snail be 6 grins. The minimum test levelfor components weighing more than 22.7 kg (50 lb) shall beevaluated on an individual basis. The test duration in each ofthe three orthogonal axes shall equal or exceed the expected‘flight exposure time, but shall not be less than 1 minute peraxis. Where insufficient time is available at the full testlevel to test all modes, extended testing at a level 6 dB lowershall be conducted as necessary to complete functional testing.

I 7.3.4.4 Supplementary Requirements. Same as 6.4.5.4.

~7.3.5 Acoustic Test, Component Acceptance

7.3.5.1 Purpose. This test detects material andworkmanship defects that might not be detected in a static testcondition.

I 7.3.5.2 Test Description. Same as 6.4.6.2.

7.3.5.3 Test Levels and Duration. The component shall beexposed to sound pressure levels equal to the maximum predictedlevels as defined in 3.21, but not less than 138 dB overall. ●The duration shall equal or exceed the expected flight exposuretime to the maximum predicted environment but shall not be less

than 1 minute.


7.3.6 Pyro Shock Screening Test, Component Acceptance

7.3.6.1 Purpose. This test is intended to detectintermittents due to conducting particles in electroniccomponents. It is effective for particle detection since shockcan dislodge the particle and the extended duration vibrationallows the particle to move to a position where a short mayoccur and permit detection. It may also detect intermittentsdue to cracked or loose dies in electronic parts that would notbe found in normal acceptance tests.

7.3.6.2 Test Description. The screening test consists ofpyro shock followed by random vibration testing. The componentshall be mounted to a rigid fixture through the normal mountingpoints of the component. The component shall be tested in eachdirection of each of three mutually perpendicular axes (6 tests).

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The component shall be electrically energized andfunctionally sequenced through all possible operating modes,including redundancy, during the testing. Circuits should bemonitored for intermittents. A functional test shall beeconducted before and after the pyro shock test and after thevibration screening test.

7.3.6.3 Test Levels and Duration. A pyro shock test shallbe performed once in both directions of each of three mutuallyperpendicular axes at maximum predicted levels as defined in3.22. The screening vibration test level shall be 3 dB belowthe acceptance test level specified in 7.3.4.3, but no lowerthan 4.5 g rms. Each test shall consist of 5 minutes dwell testfollowed by vibration bursts consisting of 10 seconds “on” and10 seconds “off.” The number of bursts shall be such that allcircuits are monitored at least 10 times during the bur8tsequence with a minimum number of 20 bursts.

7.3.6.4 Supplementary Requirements. This test has provedto be very perceptive for detecting intermittents in guidancecomparents and should be considered where the component iscritical to mission success. Rether than basing the screeningvibration test on the expected flight vibration sPectrum shaPe,a special purpose spectrum may be developed. It is believed bysome researchers that the screening vibration spectrum imposedon the component should cause circuit boards within to vibrateat a level of 5 - 10 g cms with a flat spectrum from 10 - 450Hz. The required input spectrum may be defined from results ofdevelopmental vibration survey tests or analytic responsepredictions.

The screening test incurs exposure of the component tovibration in addition to that experienced in the acceptancevibration test. The additional exposure is defined by thescreening spectrum based on the expected flight shape, or by thespecially developed spectrum if that option is selected. Ineither case, the component qualification vibration levels anddurations specified in this standard shall be assessed to assurethat they provide a sufficient margin to safely conduct thescreening test.

7.3.7 Pressure Test, Component Acceptance

7.3.7.1 Pur ose.+“

This test detects material andworkmanship de ects which could result in failure of thepressure vessel or valves in usage.

7.3.7.2 Test Description. ThiS test iS the same asdescribed in 6.4.10.2a and o, except that onlY one cYcle shall

71



be required, and test at elevated temperature is optional.

7.3.7.3 Test Levels. Same as 6.4.10.3.●

7.3.7.4 Supp lementary Requirements. Applicable safetystandards shall be followed in conducting all tests.

7.3.8 Leakage Test, Component Acceptance

7.3.8.1 Purpose. This test demonstrates the capabilitypressurized components to meet the leakage rate requirementsspecified in the component specifications.

of

7.3.8.2 Test Description and Alternatives. The componentleak checks shall be made before and after exposure to eachenvironmental acceptance test. The test method employed shallhave sensitivity and accuracy consistent with the componentsspecified maximum allowable leak rate. One of the methods givenin 6.4.11.2 shall be used.

7.3.8.3 Test Levels and Duration. Same as 6.4.11.3.

7.3.9 Burn-In Test, Component Acceptance

7.3.9.1 Purpose. The purpose of the burn-in test shall beto detect material and workmanship defects which occur early in ●the component life.

7.3.9.2 Test Description. A modified thermal cycling testshall be used to accumulate the additional operational hoursrequired for the burn-in test of electronic and electricalcomponents. While the component is operating (power on) andwhile perceptive parameters are being monitored, the temperatureof the unit shall be reduced to the specified low temperaturelevel. The unit shall be operated at tne low temperature levelfor 1 hour or longer. The unit temperature shall then beincreased to the specified high temperature level and operatedfor 1 hour or longer. The temperature shall then be reduced toambient to complete one cycle of the burn-in test. Thetransitions between low and high temperatures shall be at anaverage rate greater than 1 deg C per minute.

For valves, thrusters, and other items where the number ofCyClt?Sof operation rather than hOuKS of operation is a bettermethod to ensure detecting infant mortality failures, functionalcycling shall be conducted at ambient temperature. Forthrusters, a cycle is a hot firing which includes a start,steady-state operation, and shutdown. For hot ’firings ofthrusters utilizing hydrazine propellants, action shall be taken

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MIL-STD-1540B [USAF)I 10 OCTOBER 1982

to assure that the flight valves are thoroughly cleaned of alltraces of hydrazine following the test firings. Devices thathave extremely limited life cycles such as positive expulsiontankg are excluded from burn-in test requirements.


a. Pressure. Ambient pressure should normally beused.

b. Temperature. For cycling of electronic andelectrical components, the extreme temperaturesspecified in 7.3.3.3.b shall be used.

c. Duration. The total operating time forelectronic and electrical component burn-in shallbe 300 hours including the operating time duringthermal cycling per 7,3.3. The minimum number oftemperature cycles shall be 18 including thoseconducted during the thermal cycling acceptancetest. Additional test time beyond that requiredfor thermal cycling shall be conducted at eithermaximum or minimum temperature. The last 100hours of the component burn-in test shall be freeof failures. For valves, thrusters, and othercomponents where functional cyclic testing is abetter burn-in method, a minimum of 100 cyclesshall be conducted.

d. Functional Cuty Cycle. Functional tests shall beconducted at the start of this test to provide abaseline reference for determining if performancedegradation occurs. The functional test shall berepeated after 150 hours of operation and duringthe last 2 hours of the thermal cycling test.perceptive parameters for all circuits, includingall redundancy, shall be monitored to the maximumextent possible during the entire test sequence.On-off cycling of the electronics component shallbe conducted during the test to simulateoperational usage.

7.3.9.4 Supplementary Requirements. The reduction ofsystem level failures by burn-in at the component level has afavorable impact on costs and schedules by stabilizing thefailure rate at or near its minimum and ensuring the highest

probability of mission success.


MIL!-STD-1540B (USAF)10 OCTOBER 1982

7.4 SUBASSEMBLY LEVEL ACCEPTANCE TESTS

These tests are to be considered as optional unlessspecified otherwise in the contract. However, subassembly levelacceptance tests are often cost-effective measures for reducingor avoiding failures in higher level tests and possibly inorbital operations. Acceptance test should be conducted at thesubassembly level where this level provides a more perceptivetest than would be possible at either the part or the componentlevel. Functional or environmental acceptance tests are usuallyconducted at the subassembly level to detect material and

workmanship defects, or to measure critical parameters, thatcannot be accomplished satisfactorily at higher levels ofassembly. When these acceptance tests are planned on

subassemblies, they may be tested to similar requirements as

components, or more stringent requirements for stress screening

may be used.



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8.1

SECTION 8

FLIGHT USE OF QUALIFICATION EQUIPMENT

Qualification tests are conducted to demonstrate that thedesign, manufacturing, and assembly have resulted in hardwareconforming to specification requirements. The qualificationtests required by this document incorporate the environmentaldesign margin into the test levels to assure that flight unitswill meet the operational requirements for their service life.The vibration tests, acoustic tests, and thermal tests producecyclic stresses that can encroach on the fatigue margins ofinterconnect wiring, solder joints, structural members andsimilar items in the qualification test units. If equipmentthat has been subjected to qualification testing is planned forsubsequent flight use, it is possible that the remaining fatiguemargins are so low as to present a high risk of failure duringflight. This is primarily due to the use of high test levelsand long test durations during the baseline qualificationtests. Therefore, the actual vehicle used for the 6.2 vehiclequalification tests or the components used for the 6.4 componentqualification tests may not be suitable for subsequent flight.

Nevertheleas, initial program costs and scheduleconstraints may force the consideration of ways to make unitsused for qualification testing acceptable for flight. It shouldbe recognized that the use of qualification items for flightalways presents a higher risk than the use of standardacceptance tested items for flight. This risk may be reduced byvarious strategies such as reducing qualification test levelsand durations to reduce the encroachment on fatigue and wearoutmargins. The strategy used should be based upon specificprogram considerations. One method ha8 been to replace allcomponents on the qualification vehicle with “new” componentsthat have passed component acceptance tests (see 8.3). Anotherway was to lower the space vehicle qualification test levels andtest duration to avoid excessive encroachment on margins (see8.2). On some programs, one or more qualification componentshave been used as flight components (see 8.1). In such caseswhere program considerations are overriding, the contract maydirect, or the contracting officer may approve, the use Ofaualificatian units for fliaht. Some of the etrateqies that

been used are presente~ in the following exempies.

USE OF THE QUALIFICATION COMW2NEN’iS FOR FLIGHT

When the qualification components are planned for flight

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use, the component qualification test program shall be modifiedfrom that specified in Section 6 to reduce cyclic stresslevels. In addition, the component qualification testing shallbe conducted on flight spares so that flight use is delayed orpossibly never required. The flight space vehicle in whichthese qualification components are installed shall be acceptancetested in accordance with the requirements of 7.1. This spacevehicle qualification would be based on the requirements of 6.2.

8.1.1 Component Qualification Tests. When the componentqualification tests are conducted on a component intended for~ubsequent flight, the component acceptanc& tests required bythis standard are waived, except for the burn-in acceptance testof 7.3.9, and only the qualification test baseline specified in6.4 is required with the following exceptions:

a.

b.

c.

d.

e.

f.

For the component thermal vacuum test (6.4.2),the temperature extremes shall be 5 deg C beyondthe minimum and maximum predicted temperatures.

For the component thermal cycling test (6.4.3),the ‘temperature cycles shall be conducted at 5deg C beyond the acceptance temperature extremes(7.3.3.b)

For the component vibration qualification test(6.4.5), the test level shall be 3 dB greaterthan the maximum predicted level but not lessthan 9 grins.

For the component acoustic qualification, test(6.4.6), the test level shall be 3 dB greaterthan the maximum predicted level but not lessthan 141 dB overall.

For the component pyro shock test (6.4.7), theshock spectrum shall be 3 dB greater than themaximum predicted level.

For tbe componenc pressure test (6.4.10) only.proof pressure tests per 6.4.10.3 a and b shallbe conducted.

8.1.2 Component Certification for Flight. Upon completionof the modified qualification test program, the component testhistory shall be-reviewed for excessive test time and potentialfatigue type failures to determine if the unit is acceptable forflight or if refurbishment is required. Mission and safetycritical qualification components should not be used for flight

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in systems where a redundant compone;t is not provided.

8.2 USE OF THE FLIGHT VEHICLE FOR SPACE VEHICLE LEVELQUALIFICATION

When the flight vehicle is also used for the vehicle levelqualification tests, the space vehicle qualification test levelsand durations shall be reduced as defined in 8.2.1. Thecomponents installed in this flight vehicles shall be acceptancetested in accordance with the requirements of 7.3. Thecomponent qualifications would be based on the requirements of6.4.

8.2.1 Space Vehicle Qualification TestS. If the apacevehicle qualification tests are to be combined with the flightvehicle acceptance tests, the space vehicle level acceptancetests required by this standard are weived and only thequalification test baseline in 6.2 is required with thefollowing exceptions:

a.

b.

c.

d.

For the space vehicle acoustic qualification test(6.2.3), the test level shall be 3 dB greaterthan the maximum predicted level but not lessthan 141 dB overall. The duration of the teStshall be the same as for the space vehicleacouatic acceptance test (7.1.3.3).

For the apace vehicle vibration qualificationtest (6.2.4), the test levels shall producevibration responses in the equi~ent which are 3dB greater than the maximum predicted level. Theduration of the test shall be the same aa for thespace vehicle vibration acceptance teat (7.1.4.3). .

For the space vehicle thermal vacuumqualification test (6.2.7), the number ofhot-cold cycles shall be four and the temperatureextremes shall be 5 deg C beyond the minimum andmaximum predicted temperaturea.

If the optional apace vehicle thermal cyclingtest (6.2.9) is adopted as baseline, the minimumspace vehicle temperature range shall be 60 degc. The teat should include 15 percent morethermal cycles than the space vehicle thermalcycling acceptance test, (7.1.8.3).

8.2.2 Space Vehicle Certification for Flight. Upon

@

completion of the modified space vehicle qualification test

77


MIL-STD-1540B (USAF)10 OC’It3BER1982

program, the vehicle test history shall be reviewed forexcessive test time and potential fatigue type failure todetermine if the vehicle is acceptable for flight or ifrefurbishment is required. If significant modifications areincorporated or numerous components are refurbished or replacedwith new components subsequent to qualification testing, thespace vehicle acceptance baseline specified in 7.1 shall berequired prior to launch certification.

8.3 USE OF THE QUALIFICATION VEHICLE FOR FLIGHT

When the space vehicle used for vehicle level qualificationtesting of 6.2 is planned for subsequent flight use, allcomponents shall be replaced with “new” components that havepassed the component acceptance tests. The space vehicle iscertified for flight when it satisfactorily completes thevehicle level acceptance tests of Section 7.

8.4 OTHER ‘,

Various combinations of strategy may be considereddepending on specific program considerations and the degreerisk deemed acceptable. For example, method 8.1.1 may becombined with a vehicle qualified-at reduced levels parwith the qualification vehicle per 8.3. In such “cases,provisions of both methods apply and the resultant riskincreased appropriately.

8.2.

of

1 orthewould be ●

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9.1

SECTION 9

PRELAUNCH VALIDATION TESTS

GENERAL REQUIREMENTS

Prelaunch validation testing is accomplished at the factoryand at the launch base, with the objective of demonstratinglaunch system readiness. The test series first establishesspace vehicle baseline data in the factory preshipment test of7.1.1.3. At the launch base, these tests verify that no changeshave occurred in space vehicle parameters as a result ofhandling and transportation to the launch base. Further, aprelaunch validation test of the launch system is performed toverify that interfaces between the space vehicle and Launchvehicle, and between these vehicles and the launch facilities,are within specified limits. ‘lbthe greatest extent possible,the tests are to exercise all systems through every operationalmode, in order to ensure that all mission requirements can besatisfied.

oThe space vehicle may be delivered as a complete vehicle or

it may be assembled at the launch base with orbital injectionsubsystems, payload subsystems, propulsion subsystems, and otherspace vehicle subsystems being delivered separately. Al1factory test acceptance data shall accompany delivered flighthardware. In any case, the total launch system is firstassembled at tne launch site. The prelaunch validation testsare unique for each program in the extent of the operationsnecessary to ensure that all interfaces are properly tested.For programs which ship a complete space vehicle to the launchsite, theSe tests primarily confirm vehicle performance, checkfor transportation damage, and demonstrate interfacecompatibility.

9.2 PREIJNJNCHVALIDATION TEST FL13W

The test flow shall follow a progressive growth pattern toensure proper operation of each space vehicle element prior toprogressing to a higher level of assembly and test. In general,tests should follow the launch base buildup cycle. Aasuccessive systems or subsystems are verified, assembly proceedsto the completed space vehicle. Following testing of the spacevehicle and its interfaces with the launch vehicle, thesevehicles are electrically and mechanically mated and integratedinto the launch system. Space vehicles employing a recoverable

alaunch vehicle shall utilize a launch vehicle simulator to

79

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perform mechanical and electrical interface tests prior tointegration with the launch vehicle. Following integration ofthe space vehicle and launch vehicle, functional tests of thespace vehicle shall be conducted to ensure its proper operationfollowing the handling operations involved in mating thevehicles. Space vehicle cleanliness shall be monitored by useof witness plates.

9.3 PRELAUNCH VALIDATION TEST CONFIGURATION

During each test, the space vehicle should be in flightconfiguration to the maximum extent possible, consistent withsafety, control, and monitoring requirements. For programsutilizing a recoverable launch vehicle, the test configurationshall include any airborne support equipment required for thelaunch, ascent, and space vehicle deployment phases. Thisequipment shall be mechanically and electrically mated to thespace vehicle in its launch configuration. Whenever possible,ground support equipment should have a floating point groundscheme that is connected to the flight vehicle single pointground. Isolation resistance tests shall be run to verify thecorrect grounding scheme prior to connection to the flightvehicle. This reduces the possibility of ground equipmentinterference with vehicle performance. All ground equipmentshall be validated prior to being connected to any flighthardware, to preclude the possibility of faulty ground equipment ●causing damage to the flight hardware or inducing ambiguous or

invalid data. Test provisions shall be made to verify integrityof circuits into which flight jumpers, arm plugs, or enableplugs have been inserted.

9.4 PRELAUNCH VALIDATION TEST DESCRIPTIONS

The prelaunch validation tests shall exercise anddemonstrate satisfactory operation of the space vehicle throughall of its mission phases, to the maximum extent practical.Test data shall be compared to corresponding data obtained infactory tests to identify trends which indicate performancedegradation within specification limits. Each test procedureused shall include test limits and success criteria sufficientto permit a rapid determination as to whether or not Processing

I and integration of the vehicle should continue. Howeier, the -final acceptance or rejection decision, in most tests, dependsupon the results of post-test data analysis.

9.4.1 Functional Test. Electrical functional tests shallbe conducted that duplicate, as nearly as possible, the factoryfunctional tests of 7.1.1.2. Mechanical tests for leakage,valve and mechanism operability, and fairing clearance shall be

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conducted.

9.4.1.1 Simulators. Simulation devices shall be carefullycontrolled and shall be Dermitted onlv when there is no feasiblealternative for conducting the test. ‘When it is necessary to

employ simulators in the conduct of prelaunch validation testsof the space vehicle, the interfaces disconnected in thesubsequent replacement of the simulators with flight hardwareshall be revalidated. Simulators shall be used for thevalidation of ground support equipment prior to connecting it toflight hardware.

9.4.1.2 Explosive Circuits. When explosive circuits areinvolved, approved simulation devices shall be used whereappropriate. Before connection of pyro devices to theirrespective circuits, line continuity checks shall be made forthe presence of the “Firen signal at the squib connection whencommanded. A line continuity stray voltage check shall be madeimmediately prior to the connection of any pyro device, and thischeck shall be repeated whenever that connection is opened andprior to reconnection.

9.4.2 Propulsion System Leak and Functional ‘Ikst. Afunctional test of the space vehicle propulsion subsystem shallbe conducted to verify, to the maximum practical extent, theproper operation of all components. Propulsion system leakagerates shall be verified to be within allowable limits.

9.4.3 Integrated System Tests. Total launch systemreadiness shall be demonstrated tnrouqh an integrated, fullyassembled launch systems test prior t: flight. ‘This test shallinclude an evaluation of radio frequency (rf) interferencebetween system elements, electrical power interfaces, and thecommand and control subsystems. On a new space vehicle designor a significant design change to the telemetry, trackingt orreceiving subsystem of an existing space vehicle, a test shallbe run on the first vehicle to ensure nominal operation and thatpyrotechnics (simulators) do not fire when the vehicle issubjected to the worst case range electromagnetic interferenceenvironment.

9.4.4 Compatibility Test, On-Orbit System

the &&4&h~m”the on-orbit command and control network.This test validates the compatibility of

For the purpose of establishing this testing baseline, it isassumed that the on-orbit command and control network is (oroperationally interfaces with) the Air Force Satellite ControlFacility (AFSCF). This test demonstrates the ability of the



space vehicle, when in orbit, to properly respond to the AFSCFhardware, software, and operations team as specified in theAFSCF Orbital Support Plan. For programs which have a dedicatedground station, compatibility tests shall also be performed withthe dedicated ground station.

9.4.4.2 Test Description. Facilities to perform on-orbitsystem compatibility tests exist at the Western Test Range (WTR)and the Eastern Test Range (ETR). At both locations, the AFSCFcan command the space vehicle and process telemetry from thespace vehicle as well as perform tracking and ranging, thusverifyingtelemetry

the rf compatibility, the command software, and themodes. The tests include the following:

a. Verification of rf, analog, and digitalcompatibility of command, telemetry, and trackinglinks.

b. Verification of AFSCF capability to control thespace vehicle using single, block, unsecure, andsecure commands as required for on-orbit support.

c. Verification of AFSCF capability to process,display, and record space vehicle telemetry linkor links as required for on-orbit support.

d. Verification of AFSCF capability to track the ●space vehicle using angle, doppler, and rangetracking as required for on-orbit support.

9.4.4.3 Supplementary Requirements. This test should berun as soon as feasible after the space vehicle arrives at thelaunch base. The test is made with every space vehicle toverify system interface compatibility. The test shall be runusing the software model versions that are integrated into theoperational on-orbit software of the space vehicle under test.A preliminary compatibility test may be run prior to the arrivalof the space vehicle at the launch base by the use of prototypesubsystems, components, or simulators as required to prove theinterface. Preliminary compatibility tests may be run usingpreliminary software. Normally, a preliminary compatibilitytest is run once for each series of space vehicles to checkdesign compatibility, and is conducted well in advance of thefirst launch to permit orderly correction of hardware, software,and procedures as required. Changes in the interface from thosetested in the preliminary test shall be checked by thecompatibility tests conducted just prior to launch.

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The following Data Item Descriptions are among those mostfrequently used in the Contract Data Requirements LiSt (DD Form1423) to establish detail requirements for the preparation oftest plans, procedures, and reports.

I

.

DI-T-3701DI-T-3702DI-T-3703

DI-T-3704DI-T-3705DI-T-3707DI-T-3708DI-T-3714DI-T-3716DI-T-3717

DI-T-3718

CustodiansAir Force - 19

System Test PlanContractor Test Plans/ProceduresComputed Program Coinfiguration Item. (CPCI) TeStPlan/Procedures (Computer Programs)Electromagnetic Compatibility Test PlanStructural Test PlansGeneral Test Plan/proceduresEnvironmental Test Plans/ProceduresAcceptance Test ProceduresFinal Test ReportComputer Program Configuration Item (CPCI)Development Test and Evaluation ‘restReport

Test Reports General

Preparing ActivityAir Force - 19

(Project No. 181O-FO15)Document 0002b/Arch 0001b


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PO BOX92960WorldwayPostalCenterLoe AngeleeCA 90009

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