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Results are presented of the failure investigation conducted on resistors produced by the Vamistor Divison, Wagner Electric Corporation. This failure investigation included; failure analyses, chemical and metallurgical analyses, failure mechanism studies, seal leak analyses, and nondestructive stress tests. The data, information, conclusions, and recommendation can be helpful in assessing current usage of these resistors.
The S190B Earth Terrain Camera (ETC) operated acceptably for all of its scheduled EREP passes throughout the SL2 mission. The crew reported no problems in unstowing the camera, changing filters, installing the ETC window in the SAL, or installing the camera onto the window. The ETC was operated for a total of seven times with no failures. The clock on the ETC was checked on DOY 170 (June 19, 1973) and was found to be 30 min. and 58 sec. slower than GMT. The change in time was expected since a similar circumstance was experienced during ETC qualification testing for launch vibration. A leak existed in the seal of the spare magazine to the camera vacuum interface. For EREP passes 08 and 10, black-and-white film EK 3414 (roll no. 82) was installed in this spare magazine. Since there was an audible hiss, the vacuum hose was not connected to the camera. This caused the vacuum platen to be inoperable, resulting in some degradation in resolution for this roll of film. The vegetation of the South American jungle areas proved to be much darker than vegetation found in the United States, and was consequently about 1/2 stop underexposed in all cases.
Simulated-service cycle differs from conventional proof-test cycle in that specimen is subjected to sine-wave variation of stress instead of prolonged steady stress. Simulated-service testing is performed at low and high temperatures as well as at room temperatures. (Flaws initiated by severe stresses of proof testing can later develop into leaks during service, thus effectively reducing useful service life.) Method may be useful for commercial thin-wall vessels.
The design and test results for two types of pulsed gas valves are presented. The valves, a piezo valve and a solenoid actuated valve, must have exceedingly long lifetime to support gas-fed pulsed electric thruster operation for missions of interest. The performance of both valves was tested, with both demonstrating the capability to throttle the gas flow rate while maintaining low leakage levels below 10(exp -3) sccs of He at the beginning of valve lifetime. The piezo valve varies the flow rate by changing the amount that the valve is open, which is a function of applied voltage. This valve demonstrated continuous throttlability from 0-10 mL/s, with opening and closing times of 100 microsecond or less. The solenoid actuated valve flow rate changes as a function of the inlet gas pressure, with demonstrated flow rates in these tests from 2.7-11 mL per second. The valve response time is slower than the piezo valve, opening in 1-2 ms and closing in several ms. The solenoid actuated valve was tested to one million cycles, with the valve performance remaining relatively unchanged throughout the test. Galling of the sliding plunger caused the valve to bind and fail just after one million cycles, but at this point in the test the valve sealing surface leak rate still appeared to be well below the maximum target leak rake of 1×10(exp -3) sccs of He.
Damage tolerance is an important consideration in space structures applications. The intent of damage tolerance is to demonstrate that the structure is robust to the presence of flaws over the service life of the space vehicle. Damage tolerance requirements in documents such as those in NASA, AIAA, and ISO can be challenging to implement. The intent of this paper is three-fold: (1) Provide best practices in the application of damage tolerance requirements for space applications, (2) Provide lessons learned for each of class of hardware, (3) Provide examples of challenging situations encountered in the application of damage tolerance. Examples cover additive manufacturing, composite overwrapped pressure vessels, pressurized structures, liquid rocket engines, thermal protection systems, and other classes of hardware. The philosophy and limitations of leak before burst and proof test logic are also discussed. Finally, a discussion on elastic-plastic fracture mechanics with comparisons to test data will be presented.
This paper focuses on the flight results of the Cryogenic Two-Phase Flight Experiment (CRYOTP), which was a Hitchhiker based experiment that flew on the space shuttle Columbia in March of 1994 (STS-62). CRYOTP tested two new technologies for advanced cryogenic thermal control; the Space Heat Pipe (SHP), which was a constant conductance cryogenic heat pipe, and the Brilliant Eyes Thermal Storage Unit (BETSU), which was a cryogenic phase-change thermal storage device. These two devices were tested independently during the mission. Analysis of the flight data indicated that the SHP was unable to start in either of two attempts, for reasons related to the fluid charge, parasitic heat leaks, and cryocooler capacity. The BETSU test article was successfully operated with more than 250 hours of on-orbit testing including several cooldown cycles and 56 freeze/thaw cycles. Some degradation was observed with the five tactical cryocoolers used as thermal sinks, and one of the cryocoolers failed completely after 331 hours of operation. Post-flight analysis indicated that this problem was most likely due to failure of an electrical controller internal to the unit.
Evaluation tests are reported for 20 cells to be put into the life cycle program. Leak, capacity, internal resistance, internal short, and overcharge tests are described.
The J-2X Engine (built by Pratt & Whitney Rocketdyne,) in the Upper Stage of the Ares I Crew Launch Vehicle, will only start within a certain range of temperature and pressure for Liquid Hydrogen and Liquid Oxygen propellants. The purpose of the Simulink Integrated Performance Analysis Model is to verify that in all reasonable conditions the temperature and pressure of the propellants are within the required J-2X engine start boxes. In order to run the simulation, test variables must be entered at all reasonable values of parameters such as heat leak and mass flow rate. To make this testing process as efficient as possible in order to save the maximum amount of time and money, and to show that the J-2X engine will start when it is required to do so, a graphical user interface (GUI) was created to allow the input of values to be used as parameters in the Simulink Model, without opening or altering the contents of the model. The GUI must allow for test data to come from Microsoft Excel files, allow those values to be edited before testing, place those values into the Simulink Model, and get the output from the Simulink Model. The GUI was built using MATLAB, and will run the Simulink simulation when the Simulate option is activated. After running the simulation, the GUI will construct a new Microsoft Excel file, as well as a MATLAB matrix file, using the output values for each test of the simulation so that they may graphed and compared to other values.
Activities described include: (1) determining the applicability of an ultrahigh-strength graphite fiber to composite pressure vessels; (2) defining the fatigue performance of thin-titanium-lined, high-strength graphite/epoxy pressure vessel; (3) selecting epoxy resin systems suitable for filament winding; (4) studying the fatigue life potential of Kevlar 49/epoxy pressure vessels; and (5) developing polymer liners for composite pressure vessels. Kevlar 49/epoxy and graphite fiber/epoxy pressure vessels, 10.2 cm in diameter, some with aluminum liners and some with alternation layers of rubber and polymer were fabricated. To determine liner performance, vessels were subjected to gas permeation tests, fatigue cycling, and burst tests, measuring composite performance, fatigue life, and leak rates. Both the metal and the rubber/polymer liner performed well. Proportionately larger pressure vessels (20.3 and 38 cm in diameter) were made and subjected to the same tests. In these larger vessels, line leakage problems with both liners developed the causes of the leaks were identified and some solutions to such liner problems are recommended.
Inductive pulsed plasma thrusters are spacecraft propulsion devices in which energy is capacitively stored and then discharged through an inductive coil. While these devices have shown promise for operation at high efficiency on a range of propellants, many technical issues remain before they can be used in flight applications. A conical theta-pinch thruster geometry was fabricated and tested to investigate potential improvements in propellant utilization relative to more common, flat-plate planar coil designs. A capacitor charging system is used to permit repetitive discharging of thrusters at multiple cycles per second, with successful testing accomplished at a repetition-rate of 5 Hz at power levels of 0.9, 1.6, and 2.5 kW. The conical theta-pinch thruster geometry was tested at cone angles of 20deg, 38deg, and 60deg, with single-pulse operation at 500 J/pulse and repetitionrate operation with the 38deg model quantified through direct thrust measurement using a hanging pendulum thrust stand. A long-lifetime valve was designed and fabricated, and initial testing was performed to measure the valve response and quantify the leak rate at beginning-of-life. Subscale design and testing of a capacitor charging system required for operation on a spacecraft is reported, providing insights into the types of components needed in the circuit topology employed. On a spacecraft, this system would accept as input a lower voltage from the spacecraft DC bus and boost the output to the high voltage required to charge the capacitors of the thruster.
Discovery of leakage on several International Space Station U.S. Laboratory Module ammonia system quick disconnects (QDs) led to the need for a process to quantify total leakage without removing the QDs from the system. An innovative solution was proposed allowing quantitative leak rate measurement at ambient external pressure without QD removal. The method utilizes a helium mass spectrometer configured in the detector probe mode to determine helium leak rates inside a containment hood installed on the test component. The method was validated through extensive developmental testing. Test results showed the method was viable, accurate and repeatable for a wide range of leak rates. The accumulation method has been accepted by NASA and is currently being used by Boeing Huntsville, Boeing Kennedy Space Center and Boeing Johnson Space Center to test welds and valves and will be used by Alenia to test the Cupola. The method has been used in place of more expensive vacuum chamber testing which requires removing the test component from the system.
Thermal performance measurement of piping systems under actual field conditions is important for space launch development and commercial industry. Knowledge of the true insulating effectiveness is needed in system design, development, and research activities. A new 18-meter-long test apparatus for cryogenic pipelines has been developed. Three different pipelines, rigid or flexible, can be tested simultaneously. Critical factors in heat leak measurements include eliminating heat transfer at end connections and obtaining proper liquid saturation condition. Effects due to variations in the external ambient conditions like wind, humidity, and solar radiation must be minimized. The static method of liquid nitrogen evaporation has been demonstrated, but the apparatus can be adapted for dynamic testing with cryogens, chilled water, or other working fluids. This technology is suited for the development of an industry standard test apparatus and method. Examples of the heat transfer data from testing commercially available pipelines are given. Prototype pipelines are currently being tested and evaluated at the Cryogenics Test Laboratory of NASA Kennedy Space Center.
After a decade of construction and a year of activation the El Ultra High Pressure Propulsion Test Facility at NASA's Stennis Space Center is fully operational. The El UHP Propulsion Test Facility is a multi-cell, multi-purpose component and engine test facility . The facility is capable of delivering cryogenic propellants at low, high, and ultra high pressures with flow rates ranging from a few pounds per second up to two thousand pounds per second. Facility activation is defined as a series of tasks required to transition between completion of construction and facility operational readiness. Activating the El UHP Propulsion Test Facility involved independent system checkouts, propellant system leak checks, fluid and gas sampling, gaseous system blow downs, pressurization and vent system checkouts, valve stability testing, valve tuning cryogenic cold flows, and functional readiness tests.
Electrified aircraft are being developed to increase the efficiency and reduce the cost of operating subsonic transport aircraft. Achieving a substantial impact necessitates focusing on single- and twin-aisle aircraft which use propulsion systems with >20 MW ratings. For these aircraft, multi-MW superconducting electric machines are being developed due to their high specific power and high efficiency. Most of these electric machines employ superconductors on their rotor, thereby requiring cryogenic cooling of the rotating system. An attractive solution is to conductively cool the rotor using a cryocooler that rotates with the rotor. NASA has been developing a 50 K pulse tube cryocooler for its High Efficiency Megawatt Motor (HEMM) that can operate while rotating at up to 6,800 rpm. The first non-rotating tests of the fully assembled HEMM cryocooler have been completed, and initial testing has yielded a no-load temperature of 112 K. A small helium leak into the vacuum system has been identified as the major bottleneck to performance, and based on measurements with the motor turned off there is an observed vacuum-side heat leak of approximately 30 W at 115 K. Immediate next steps for testing are to locate and mitigate the apparent helium leaks into the vacuum chamber. Thus far, compressor performance and the apparent heat lift after correcting for the heat leak are on par with expectations.
The James Webb Space Telescope will undergo a full system test in the cryogenic vacuum chamber A at the Johnson Spaceflight Center in order to verify the overall performance of the combined telescope and instrument suite. This will be the largest and most extensive cryogenic test ever undertaken. Early in the test system development, it was determined that precise position measurements of the overall hardware would enhance the test results. Various concepts were considered before selecting photogrammetry for this metrology. Photogrammetry has been used in space systems for decades, however cryogenic use combined with the size and the optical/thermal sensitivity of JWST creates a unique set of implementation challenges. This paper provides an overview of the JWST photogrammetric system and mitigation strategies for three key engineering design challenges: 1) the thermal design of the viewing windows to prevent excessive heat leak and stray light to the test article 2) cost effective motors and mechanisms to provide the angle diversity required, and 3) camera-flash life and reliability sufficient for inaccessible use during the number and duration of the cryogenic tests.
There is a limited amount of hypervelocity impact (HVI) data on pressurized composite overwrapped pressure vessels (COPV). In recent years, NASA has performed HVI tests to characterize impact conditions resulting in either leak or burst of the COPVs representative of spacecraft hardware. This paper reports on the results of 40 tests that have been conducted on several types of COPV configurations, pressurized by inert gas to near the vessels rated maximum expected operating pressure (MEOP). These tests were used to better understand COPV response under HVI conditions and develop ballistic limit equations (BLE) related to these tests.
The Data Acquisition and Control Systems (DACS) Laboratory is a facility at Stennis Space Center that provides an off test-stand capability to develop data-acquisition and control systems for rocket-engine test stands. It is also used to train new employees in state-of-the-art systems, and provides a controlled environment for troubleshooting existing systems, as well as the ability to evaluate the application of new technologies and process improvements. With the SSC propulsion testing schedules, without the DACS Laboratory, it would have been necessary to perform most of the development work on actual test systems, thereby subjecting both the rocket-engine testing and development programs to substantial interference in the form of delays, restrictions on modifications of equipment, and potentially compromising software configuration control. The DACS Laboratory contains a versatile assortment of computer hardware and software, digital and analog electronic control and data-acquisition equipment, and standard electronic bench test equipment and tools. Recently completed Control System development and software verification projects include support to the joint NASA/Air Force Integrated Powerhead Demonstration (IPD) LOX & LH2 PreBurner and Turbopump ground testing programs. In other recent activities, the DACS Laboratory equipment and expertise have supported the off-stand operation of high-pressure control valves to correct valve leak problems prior to installation on the test stand. Future plans include expanding the Laboratory's capabilities to provide cryogenic control valve characterization prior to installation, thereby reducing test stand activation time.
A prototype integrated liquid oxygen/liquid methane propulsion system was hot-fire tested at a variety of simulated altitude and thermal conditions in the NASA Glenn Research Center Plum Brook Station In-Space Propulsion Thermal Vacuum Chamber (formerly B2). This test campaign served two purposes: 1) Characterize the performance of the Plum Brook facility in vacuum accumulator mode and 2) Collect the unique data set of an integrated LOX/Methane propulsion system operating in high altitude and thermal vacuum environments (a first). Data from this propulsion system prototype could inform the design of future spacecraft in-space propulsion systems, including landers. The test vehicle for this campaign was the Integrated Cryogenic Propulsion Test Article (ICPTA), which was constructed for this project using assets from the former Morpheus Project rebuilt and outfitted with additional new hardware. The ICPTA utilizes one 2,800 lbf main engine, two 28 lbf and two 7 lbf reaction control engines mounted in two pods, four 48-inch propellant tanks (two each for liquid oxygen and liquid methane), and a cold helium system for propellant tank pressurization. Several hundred sensors on the ICPTA and many more in the test cell collected data to characterize the operation of the vehicle and facility. Multiple notable experiments were performed during this test campaign, many for the first time, including pressure-fed cryogenic reaction control system characterization over a wide range of conditions, coil-on-plug ignition system demonstration at the vehicle level, integrated main engine/RCS operation, and a non-intrusive propellant mass gauging system. The test data includes water-hammer and thermal heat leak data critical to validating models for use in future vehicle design activities. This successful test campaign demonstrated the performance of the updated Plum Brook In-Space Propulsion thermal vacuum chamber and incrementally advanced the state of LOX/Methane propulsion technology through numerous system-level and subsystem experiments.