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The test plan outlines the acceptance testing including sequence of testing, responsibilities, and scope of tests for the bench test equipment for the ultraviolet spectrometer experiment. The BTE is classified as mission support equipment.
A compilation of engineering design tests which were conducted in support of the Energetic Particle Satellite S-3, S-3A, and S-3b programs. The purpose for conducting the tests was to determine the adequacy and reliability of the Energetic Particles Series of satellites designs. The various tests consisted of: (1) moments of inertia, (2) functional reliability, (3) component and structural integrity, (4) initiators and explosives tests, and (5) acceptance tests.
The space shuttle solid rocket boosters (SRB's) will be jettisoned to impact in the ocean within a 200-mile radius of the launch site. Tests were conducted at Long Beach, California, using a 12-inch diameter Titan 3C model to simulate the full-scale characteristics of the prototype SRB during retrieval operations. The objectives of the towing tests were to investigate and assess the following: (1) a floating and towing characteristics of the SRB; (2) need for plugging the SRB nozzle prior to tow; (3) attach point locations on the SRB; (4) effects of varying the SRB configuration; (5) towing hardware; and (6) difficulty of attaching a tow line to the SRB in the open sea. The model was towed in various sea states using four different types and varying lengths of tow line at various speeds. Three attach point locations were tested. Test data was recorded on magnetic tape for the tow line loads and for model pitch, roll, and yaw characteristics and was reduced by computer to tabular printouts and X-Y plots. Profile and movie photography provided documentary test data.
Experimental aerodynamic investigations were conducted on a .0405 scale representation of the space shuttle orbiter in a 7.75 x 11 foot low speed wind tunnel during the time period March 21, to April 17, 1973. The primary test objectives were to investigate both the aerodynamic and propulsion effects of various air breathing engine systems in free air and in the presence of the ground. The free air portion of this test investigated the aerodynamic effects of engine nacelle number, nacelle grouping, and nacelle location. For this testing the model was sting mounted on a six component internal strain gage balance entering through the model base. The ground plane portion of the aerodynamic test investigated the same nacelle effects at ground plane locations of full scale W.P. = 239.9, 209.3, 158.9, 108.5, and 7.78 in. At the conclusion of the aerodynamic test period the propulsion effects of various nacelle locations and freestream orientations in the presence of the ground were investigated.
The Linear Test Bed program was to design, fabricate, and evaluation test an advanced aerospike test bed which employed the segmented combustor concept. The system is designated as a linear aerospike system and consists of a thrust chamber assembly, a power package, and a thrust frame. It was designed as an experimental system to demonstrate the feasibility of the linear aerospike-segmented combustor concept. The overall dimensions are 120 inches long by 120 inches wide by 96 inches in height. The propellants are liquid oxygen/liquid hydrogen. The system was designed to operate at 1200-psia chamber pressure, at a mixture ratio of 5.5. At the design conditions, the sea level thrust is 200,000 pounds. The complete program including concept selection, design, fabrication, component test, system test, supporting analysis and posttest hardware inspection is described.
Silver/Teflon thermal control coatings have been tested on a modular radiator system projected for use on the space shuttle. Seven candidate adhesives have been evaluated in a thermal vacuum test on radiator panels similar to the anticipated flight hardware configuration. Several classes of adhesives based on polyester, silicone, and urethane resin systems were tested. These included contact adhesives, heat cured adhesives, heat and pressure cured adhesives, pressure sensitive adhesives, and two part paint on or spray on adhesives. The coatings attached with four of the adhesives, two silicones and two urethanes, had no changes develop during the thermal vacuum test. The two silicone adhesives, both of which were applied to the silver/Teflon as transfer laminates to form a tape, offered the most promise based on application process and thermal performance. Each of the successful silicone adhesives required a heat and pressure cure to adhere during the cryogenic temperature excursion of the thermal-vacuum test.
A transfer function vibration test was made on an 80 ton open hopper freight car in an effort to obtain validation data on the car's nonlinear elastic model. Test configuration, handling, test facilities, test operations, and data acquisition/reduction activities necessary to meet the conditions of test requirements are given.
Variations of nozzle performance characteristics of the model nozzles used in the Space Shuttle IA12B, IA12C, IA36 power-on launch vehicle test series are shown by comparison between experimental and analytical data. The experimental data are nozzle wall pressure distributions and schlieren photographs of the exhaust plume shapes. The exhaust plume shapes were simulated experimentally with cold flow while the analytical data were generated using a method-of-characteristics solution. Exhaust plume boundaries, boundary shockwave locations and nozzle wall pressure measurements calculated analytically agree favorably with the experimental data from the IA12C and IA36 test series. For the IA12B test series condensation was suspected in the exhaust plumes at the higher pressure ratios required to simulate the prototype plume shapes. Nozzle calibration tests for the series were conducted at pressure ratios where condensation either did not occur or if present did not produce a noticeable effect on the plume shapes. However, at the pressure ratios required in the power-on launch vehicle tests condensation probably occurs and could significantly affect the exhaust plume shapes.
Tests were conducted in a hypersonic wind tunnel using various truncated space shuttle orbiter configurations in an attempt to establish the optimum model size for other tests examining body shock-wing leading edge interference effects. The tests were conducted at Mach number 8 using the phase change paint technique. A test description, tabulated data, and tracings of isotherms made from photographs taken during the test are presented.
An arc heater shroud test technique was developed for thermostructural testing of full-scale, missile nose tip-skirt assemblies. Reentry trajectory simulation is provided by changing the arc heater operation in a prescribed manner through manual and automatic controls. Excellent test repeatability allows screening of materials by direct comparison of test results. Continuous recordings of model surface temperature, pressure, heat flux, and model internal temperatures are made at several locations during the test.
The Air Force Flight Test Center (AFFTC) flight flutter facility is described. Concepts of using a minicomputer-based time series analyzer and a modal analysis software package for flight flutter testing are examined. The results of several evaluations of the software package are given. The reasons for employing a minimum phase concept in analyzing response only signals are discussed. The use of a Laplace algorithm is shown to be effective for the modal analysis of time histories in flutter testing. Sample results from models and flight tests are provided. The limitations inherent in time series analysis methods are discussed, and the need for effective noise reduction techniques is noted. The use of digital time series analysis techniques in flutter testing is shown to be fast, accurate, and cost effective.
Failure data from 16 commercial spacecraft were analyzed to evaluate failure trends, reliability growth, and effectiveness of tests. It was shown that the test programs were highly effective in ensuring a high level of in-orbit reliability. There was only a single catastrophic problem in 44 years of in-orbit operation on 12 spacecraft. The results also indicate that in-orbit failure rates are highly correlated with unit and systems test failure rates. The data suggest that test effectiveness estimates can be used to guide the content of a test program to ensure that in-orbit reliability goals are achieved.
Results of tests OS8A and B and pertinent test and model information are presented. The test was conducted in two parts. Test OS8A was performed in the NASA/ARC unitary 11-foot section and OS8B was conducted in the NASA/ARC unitary 9 x 7 tunnel. Test objectives were to investigate thermal protection system (TPS) tile sensitivity to extreme pressure gradients and vibration and to define the TPS aerodynamic environment. The model utilized for OS8 was a full-scale representation of a portion of the space shuttle orbiter lower wing carry-through structure forward of the aft tank strut and propellant lines. Thermal protection tiles were simulated on the model. The model was instrumented with static pressure orifices, accelerometers, and dynamic pressure transducers (Kulites). A total rake was utilized to obtain boundary layer surveys.
The Definition Study of a Variable Cycle Experimental Engine (VCEE) and Associated Test Program and Test Plan, was initiated to identify the most cost effective program for a follow-on to the AST Test Bed Program. The VCEE Study defined various subscale VCE's based on different available core engine components, and a full scale VCEE utilizing current technology. The cycles were selected, preliminary design accomplished and program plans and engineering costs developed for several program options. In addition to the VCEE program plans and options, a limited effort was applied to identifying programs that could logically be accomplished on the AST Test Bed Program VCE to extend the usefulness of this test hardware. Component programs were provided that could be accomplished prior to the start of a VCEE program.
A test procedure was developed and applied to terrestrial silicon solar cells in order to determine the effect of accelerated environmental and time-temperature aging on metal contact integrity. Quantities of cells of four different manufacturers were given the contact integrity test after being subjected to accelerated stress tests that included forward bias-temperature, thermal cycle and thermal shock, power cycle, and bias-temperature humidity tests at two temperature-humidity levels. Significant effects due to certain stress tests were found for some cell types. It is concluded that cells fabricated using plated nickel/solder metallization showed significantly more serious contact integrity degradation than silver-metallized cells.
A combined experimental and analytical program to develop system electrical test procedures for the qualification of spacecraft against damage produced by space-electron-induced discharges (EID) occurring on spacecraft dielectric outer surfaces is described. The data on the response of a simple satellite model, called CAN, to electron-induced discharges is presented. The experimental results were compared to predicted behavior and to the response of the CAN to electrical injection techniques simulating blowoff and arc discharges. Also included is a review of significant results from other ground tests and the P78-2 program to form a data base from which is specified those test procedures which optimally simulate the response of spacecraft to EID. The electrical and electron spraying test data were evaluated to provide a first-cut determination of the best methods for performance of electrical excitation qualification tests from the point of view of simulation fidelity.
The Advanced Flexible Reusable Surface Insulation (AFRSI) test article was wind tunnel tested. The AFRSI was exposed to a simulated ascent airloads environment and data was obtained which could be used to support the AFRSI certification program. The AFRSI sequence of environments also included radiant heating (1500 degrees Fahrenheit) and wind/rain environments. The test article was wind/rain conditioned before each wind tunnel entry and was thermally conditioned after each wind tunnel entry. The AFRSI failed and the test was aborted before reaching the ascent environment. The AFRSI test article sequentially exposed to 50 wind/rain and 49 simulated entry thermal missions, as well as four wind tunnel entries equivalent to 40 ascent missions.
Shuttle communications and tracking systems space to space and space to ground compatibility and performance evaluations are conducted in the NASA Johnson Space Center Electronics Systems Test Laboratory (ESTL). This evaluation is accomplished through systems verification/certification tests using orbiter communications hardware in conjunction with other shuttle communications and tracking external elements to evaluate end to end system compatibility and to verify/certify that overall system performance meets program requirements before manned flight usage. In this role, the ESTL serves as a multielement major ground test facility. The ESTL capability and program concept are discussed. The system test philosophy for the complex communications channels is described in terms of the major phases. Results of space to space and space to ground systems tests are presented. Several examples of the ESTL's unique capabilities to locate and help resolve potential problems are discussed in detail.