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Characteristics of shock loads on spacecraft structures produced by pyrotechnics and explosive devices - Vol. 1
Compilation of shock loads on spacecraft structures produced by actuation of pyrotechnics and explosive devices - Vol. 4
Linear shaped charge (LSC) literature for the past 20 years is reviewed. The following topics are discussed: (1) LSC configuration; (2) LSC usage; (3) LSC induced pyroshock; (4) simulated pyrotechnic testing; (5) actual pyrotechnic testing; (6) data collection methods; (7) data analysis techniques; (8) shock reduction methods; and (9) design criteria. Although no new discoveries have been made in LSC research, charge shapes are improved to allow better cutting performance, testing instrumentation is refined, and some new explosives, for use in LSC, are formulated.
Determination of shock loads on spacecraft structures created by explosive separation of components - Vol. 3
Characteristics of shock loads induced on spacecraft structures by initiation of explosive pin pullers - Vol. 5
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Research program provides better understanding of pyrotechnic phenomenon for design purposes. Evaluating potential for damage to spacecraft by activation of pyrotechnic mechanisms, pyrotechnic-shock tests conducted on three configurations: pin pullers on orthogonal double Hopkinson bar arrangement; pin pullers on mockup of Halogen Occultation Experiment (HALOE) structure; and section of separation joint on single Hopkinson bar. Strains and accelerations measured. Strains converted to output stresses, forces, and moments. Acceleration shock-response spectra obtained for both acceleration and force signals. Results of research useful to designers in making comparison and evaluation tests before committing to costly spacecraft hardware.
Throughout the aerospace industry, large variations of 50 percent (6 dB) or more in shock response spectra (SRS) derived from pyrotechnic separation events continue to be reported from actual spaceflight data and from laboratory tests. As a result of these variations, NASA funded a research program for 1984 through 1986. The purpose of the 1984 through 1986 project was to analyze variations in pyrotechnically induced SRS and to determine if and to what degree manufacturing and assembly variables and tolerances, distance from the shock source, data acquisition instrumentation, and shock energy propagation affect the SRS. Sixty-four free-free boundary plate tests were performed. NASA funded an additional study for 1987 through 1988. This paper is a summary of the additional study. The purpose was to evaluate shock dissipation through various spacecraft structural joint types, to evaluate shock variation for various manufacturing and assembly variables on clamped boundary test plates, and to verify data correction techniques. Five clamped boundary plate tests investigated manufacturing and assembly variables and mass loading effects. Six free-free boundary plate tests investigated shock dissipation across spacecraft joint structures.
Throughout the aerospace industry, large variations of 50 percent (6 dB) or more are continually noted for linear shaped charge (LSC) generated shock response spectra (SRS) from flight data (from the exact same location on different flights) and from plate tests (side by side measurements on the same test). A research program was developed to investigate causes of these large SRS variations. A series of ball drop calibration tests to verify calibration of accelerometers and a series of plate tests to investigate charge and assembly variables were performed. The resulting data were analyzed to determine if and to what degree manufacturing and assembly variables, distance from the shock source, data acquisition instrumentation, and shock energy propagation affect the SRS. LSC variables consisted of coreload, standoff, and apex angle. The assembly variable was the torque on the LSC holder. Other variables were distance from source of accelerometers, accelerometer mounting methods, and joint effects. Results indicated that LSC variables did not affect SRS as long as the plate was severed. Accelerometers mounted on mounting blocks showed significantly lower levels above 5000 Hz. Lap joints did not affect SRS levels. The test plate was mounted in an almost free-free state; therefore, distance from the source did not affect the SRS. Several varieties and brands of accelerometers were used, and all but one demonstrated very large variations in SRS.
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Specifications for shock testing of components that will be used on the Space Shuttle vehicles require very high acceleration levels. A special shock machine was built for testing of rocket components to determine if they can meet the specified accelerations. Calibrations of transducers and methods to monitor the shock tests raised several signature-analysis questions. In this report, calibration capabilities of shock accelerometers are found to be limited to 10,000g. Equivalency of the mechanical shock test and the rocket pyrotechnic shock are examined, and two simple relationships for equivalency are proposed. Five different pulse signature-analysis techniques are tested on analytical and experimental pulse data and recommendations are made for the signature technique which most clearly identifies the magnitude of the impulse applied to the test specimen.
This paper presents the overall approach at JPL in performing spacecraft pyrotechnic shock qualification testing. Initially, the assembly shock requirements are developed early in the program based on previous spacecraft test experience and data. Pyrotechnic device development testing firings and spacecraft Development Test Model (DTM) pyro firings are then conducted to verify the adequacy of the assembly shock requirements and to determine the subsystem test firing and the subsequent system level test firing requirements. The electro-dynamic shaker, through shock synthesis techniques, is utilized to qualify the shock sensitive flight equipment with margins applied. Actual pyrotechnic device firings on spacecraft equipment or science instruments are performed when the influence of the pyros is localized and can be ignored at the system level. Full spacecraft system level shock tests, which include multiple firings of certain critical pyro devices, are conducted to verify the spacecraft design structural integrity and functions as well as to qualify hardware items which have not been previously qualified. These tests also provide a source of data from which assembly level requirements can be evaluated and compared. For example, during the Galileo program, the results demonstrated that good agreement between predicted and measured shock environments and adequate qualification of the flight spacecraft was achieved.
The technical problems encountered in the development of an advanced separation nut design are described. The nut is capable of sustaining a large preload and releasing that load with a low level of induced pyrotechnic shock, while demonstrating a tolerance for extremely high shock imposed by other pyrotechnic devices. The analysis of the separation nut was performed to acquire additional understanding of the phenomena affecting operation of the nut and to provide quantitative evaluation of design modification for aerospace applications.
Pyrotechnic shock, or pyroshock, is the transient response of a structure to loading induced by the ignition of pyrotechnic (explosive or propellant activated) devices. These devices are typically used to separate structural systems (e.g., separate a spacecraft from a launch vehicle) and deploy appendages (e.g., solar panels). Pyroshocks are characterized by high peak acceleration, high-frequency content, and short duration. Because of their high acceleration and high-frequency, pyroshocks can cause spaceflight hardware to fail. Verifying by test that spaceflight hardware can withstand the anticipated shock environment is considered essential to mission success. The Earth Observing System (EOS) AM-1 spacecraft for NASA's Mission to Planet Earth is scheduled to be launched on an Atlas IIAS vehicle in 1999, and the NASA Lewis Research Center is the launch vehicle integrator for this NASA Goddard Space Flight Center spacecraft. The EOS spacecraft was subjected to numerous ground shock tests to verify that its scientific instruments and avionics components will withstand the shock-induced vibration produced when the spacecraft separates from the launch vehicle. Shock test data from these tests represent the third largest available pyroshock database in the United States. Future spacecraft missions will directly benefit from the knowledge gained from these tests. The payload separation system used for EOS is a new system that operates by firing six separation nuts. This system was tested to verify its functional operation and to characterize the resulting shock levels. The launch vehicle contractor (Lockheed Martin Astronautics) and spacecraft contractor (Lockheed Martin Missiles & Space) completed 16 separation test firings. This resulted in an unusually large amount of pyroshock data. Typically, only one or two pyroshock test firings are performed for a spacecraft mission. Because of the size of this separation system shock database, engineers were able to perform unique statistical analyses to characterize the distribution of the test data. For example, it was proven that the shock data follow a lognormal distribution, a concept often assumed but rarely proven. The test-to-test repeatability of the shock source level was analyzed, and the effects of various test configurations and separation nut production lots were examined and quantified. Engineers investigated the change in shock level as the shock traveled from the spacecraft separation interface to the avionics components of the upper stage and analyzed the effects of the structural fidelity (simulator versus real) of the components and their weight on vibrational response. In addition, the shock attenuation with distance and across joints was quantified and compared with concepts originally generated in 1970, and the effects of separation nut preload and firing sequences effects were examined. Because of this EOS shock testing and the analyses performed at NASA Lewis, a significant amount of new information on pyroshock and its characteristics is now available to the aerospace industry. We hope that this information will help future spacecraft test planners to perform better and cheaper spacecraft separation shock tests and to better understand their test data.
The Flight Robotics Laboratory FRL successfully demonstrated the X-38 bolt retractor subsystem (BRS). The BRS design was proven safe by testing in the Pyrotechnic Shock Facility (PSI) before being demonstrated in the FRL. This Technical Memorandum describes the BRS, FRL, PSF, and interface hardware. Bolt retraction time, spacecraft simulator acceleration, and a force analysis are also presented. The purpose of the demonstration was to show the FRL capability for spacecraft separation testing using pyrotechnics. Although a formal test was not performed due to schedule and budget constraints, the data will show that the BRS is a successful design concept and the FRL is suitable for future separation tests.
Not all zero shifted acceleration data can De-trended using this technique. DC shifts, improper AC coupling, Circuit noise/EMI/EMR, Equivalent RC circuit gain response/Circuit saturation(Slew Rate Limited), fixture grounding and wiring losses can all contribute to bad shock data being recorded. Some data that is zero-shifted or exhibit large instantaneous velocity shifts is inherently bad and a retest is warranted. Clean Acceleration-Time history data can be bad upon examining the Velocity & Displacement profiles. Laser Vibrometers provide a high level of accuracy for pyrotechnic shock testing. Engineering judgment and experience will determine the validity of Shock data.
A paper describes a demonstration of the X-38 bolt-retractor system (BRS) on a spacecraft-simulating apparatus, called the Large Mobility Base, in NASA's Flight Robotics Laboratory (FRL). The BRS design was proven safe by testing in NASA's Pyrotechnic Shock Facility (PSF) before being demonstrated in the FRL. The paper describes the BRS, FRL, PSF, and interface hardware. Information on the bolt-retraction time and spacecraft-simulator acceleration, and an analysis of forces, are presented. The purpose of the demonstration was to show the capability of the FRL for testing of the use of pyrotechnics to separate stages of a spacecraft. Although a formal test was not performed because of schedule and budget constraints, the data in the report show that the BRS is a successful design concept and the FRL is suitable for future separation tests.
Pyrotechnic shock tests for Mariner Venus 67 flight acceptance