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Presentation on intern rotations. Summer 2013 I worked in cryolab and learned about calorimitry testing. Spring 2014 I helped create themalcouple arm for GODU-LH2. Summer 2014 I assisted with hardware failures on GPIM fracture mechanics testing. Summer 2017 I created a qualification test plan and test fixture to test vacuum seal-off valves.
Proposed fixture for mounting large ball bearings during tests facilitates application of variable, known radial loads, and prevent gross movement of bearing radial or axial directions. Load applied through Belleville spring washers to collar restrained by carrier ring. Load applied smoothly because collar and outer race of bearing flex in carrier ring. Inner race mounted on shaft rotated during application load.
An experiment was conducted to determine whether a sample of shorter flexible metal hoses could sustain tensile loads of up to 200 lbf and continue to meet the minimum mission requirements. The purpose of this experiment was to determine if tension loads during the processing of the hoses compromise performance. With this information, it will be decided whether the flight flex-hose manufacturer should proceed with the testing of the full-scale flex-hoses. To reduce the time and funding required for this test, a test fixture was designed and assembled in the Engineering Development & Operations (EDO) test facility. In this test fixture, two Engineering Development Unit (EDU) flex-hoses were loaded with free floating weights up to 350 lbf. A load of 200 lbf equates to the maximum expected loading with a margin of safety, thus all data recorded after 200 lbf was purely for reference. Measurements were taken using a tape measure and a custom datum measurement system to record the loaded and unloaded length of each flex-hose at various loads. Any permanent stretching beyond a 1/8th of an inch was indicative of inelastic yielding. The loading of each flex-hose was done with an initial weight 80 lbf and was increased to 100 lbf. Additional loading up to 350 lbf was done by 50 lbf increments thereafter. During the performance of the test, slippage occurred in the mounting of the flex-hoses in the test fixture. The first slippage occurred during the testing of the first flex-hose due to the collar of the flex-hose slipping within the collet of the top Kellem. As a result, the first flex-hose test was terminated early to modify the fixture. Due to the flex-hose not inelastically yielding, the test was repeated on the first flex hose. This test resulted in another instance of slippage in the upper collar of the Kellem due to tape interfering with the securing of a collet around the collar of the flex-hose. The test was then continued with one more slippage of the flex-hose within the collet of the bottom Kellem. Preventive measures were taken for future slippage, and the second hose remained secure during testing.Once the tests were concluded, the elongation of each hose was analyzed for inelastic yielding. Both flex-hoses stretched a measurable and repeatable amount under loading, however, this stretching was recovered once each hose was unloaded. As a result, both EDU flex-hoses did not experience any inelastic yielding during the tension testing. Once received, one additional flex-hose will be tested for yielding, but at the time of this paper, the recommendation is to proceed with testing of the full-scale flex-hoses at the flex-hose manufacturer.
The fabrication of two shear web test elements and three large scale shear web test components are reported. In addition, the fabrication of test fixtures for the elements and components is described. The center-loaded beam test fixtures were configured to have a test side and a dummy or permanent side. The test fixtures were fabricated from standard extruded aluminum sections and plates and were designed to be reuseable.
NASA is developing an expendable heavy lift launch vehicle capability, the Space Launch System, to support lunar and deep space exploration. To support this capability, an updated ground infrastructure is required including modifying an existing Mobile Launcher system. The Mobile Launcher is a very large heavy beam/truss steel structure designed to support the Space Launch System during its buildup and integration in the Vehicle Assembly Building, transportation from the Vehicle Assembly Building out to the launch pad, and provides the launch platform at the launch pad. The previous Saturn/Apollo and Space Shuttle programs had integrated vehicle ground vibration tests of their integrated launch vehicles performed with simulated free-free boundary conditions to experimentally anchor and validate structural and flight controls analysis models. For the Space Launch System program, the Mobile Launcher will be used as the modal test fixture for the ground vibration test of the first Space Launch System flight vehicle, Exploration Mission ? 1( now referred to as Artemis 1), programmatically referred to as the Integrated vehicle modal test. The Integrated vehicle modal test of the Exploration Mission - 1 integrated launch vehicle will have its core and second stages unfueled while mounted to the ML while inside the Vehicle Assembly Building, which is currently scheduled for the late spring or early summer of 2020. The Space Launch System program has implemented a building block approach for dynamic model validation. The modal test of the Mobile Launcher is an important part of this building block approach in supporting the integrated vehicle modal test since the Mobile Launcher will serve as a structurally dynamic test fixture whose modes will couple with the modes of the Exploration Mission ? 1 test vehicle. The Mobile Launcher modal test will further support understanding the structural dynamics of the Mobile Launcher and SLS during rollout to the launch pad, which will play a key role in better understanding and prediction of the rollout forces acting on the launch vehicle. The Mobile Launcher modal test is currently scheduled for the summer of 2019. Due to a very tight modal testing schedule, this Mobile Launcher modal pretest analysis has been performed to ensure there is a high likelihood of being able to successfully complete the modal test (i.e. identify the primary target modes) using the planned instrumentation, shakers, and excitation types. This paper will discuss this Mobile Launcher modal pretest analysis and the unique challenges faced due to the Mobile Launcher's size and weight, which are typically not faced when modal testing aerospace structures.
Flight test and theoretical aerodynamic data were obtained for a flight test fixture mounted on the underside of an F-104G aircraft. The theoretical data were generated using two codes, a two dimensional transonic code called Code H, and a three dimensional subsonic and supersonic code call wing-body. Pressure distributions generated by the codes for the flight test fixture as well as boundary layer displacement thickness generated by the two dimensional code were compared to the flight test data. The two dimensional code pressure distributions compared well except at the minimum pressure point and trailing edge. Shock locations compared well except at high transonic speeds. The three dimensional code pressure distributions compared well except at the trailing edge of the flight test fixture. The two dimensional code does not predict displacement thickness of the flight test fixture well.
NASA is developing an expendable heavy lift launch vehicle capability, the Space Launch System, to support lunar and deep space exploration. To support this capability, an updated ground infrastructure is required including modifying an existing Mobile Launcher system. The Mobile Launcher is a very large heavy beam/truss steel structure designed to support the Space Launch System during its buildup and integration in the Vehicle Assembly Building, transportation from the Vehicle Assembly Building out to the launch pad, and provides the launch platform at the launch pad. The previous Saturn/Apollo and Space Shuttle programs had integrated vehicle ground vibration tests of their integrated launch vehicles performed with simulated free-free boundary conditions to experimentally anchor and validate structural and flight controls analysis models. For the Space Launch System program, the Mobile Launcher will be used as the modal test fixture for the ground vibration test of the first Space Launch System flight vehicle, Artemis 1, programmatically referred to as the integrated vehicle modal test. The integrated vehicle modal test of the Artemis 1 integrated launch vehicle will have its core and second stages unfueled while mounted to the Mobile Launcher while inside the Vehicle Assembly Building, which is currently scheduled for the summer of 2020. The Space Launch System program has implemented a building block approach for dynamic model validation. The modal test of the Mobile Launcher is an important part of this building block approach in supporting the integrated vehicle modal test since the Mobile Launcher will serve as a structurally dynamic test fixture whose modes will couple with the modes of the Artemis 1 integrated vehicle. The Mobile Launcher modal test will further support understanding the structural dynamics of the Mobile Launcher and Space Launch System during rollout to the launch pad, which will play a key role in better understanding and prediction of the rollout forces acting on the launch vehicle. The Mobile Launcher modal test is currently scheduled for the summer of 2019. Due to a very tight modal testing schedule, this independent Mobile Launcher modal pretest analysis has been performed to ensure there is a high likelihood of successfully completing the modal test (i.e. identify the primary target modes) using the planned instrumentation, shakers, and excitation types. This paper will discuss this Mobile Launcher modal pretest analysis for its three test configurations and the unique challenges faced due to the Mobile Launcher’s size and weight, which are typically not faced when modal testing aerospace structures.
Flight test and theoretical aerodynamic data were obtained for a flight test fixture mounted on the underside of an F-104G aircraft. The theoretical data were generated using two codes: a two-dimensional transonic code called code H, and a three-dimensional subsonic and supersonic code called wing-body. Pressure distributions generated by the codes for the flight test fixture, as well as compared with the flight-measured data. The two-dimensional code pressure distributions compared well except at the minimum pressure point and the trailing edge. Shock locations compared well except at high transonic speeds. However, the two-dimensional code did not adequately predict the displacement thickness of the flight test fixture. The three-dimensional code pressure distributions compared well except at the trailing edge of the flight test fixture.
The Aerostructures Test Wing (ATW) was an apparatus used in a flight experiment during a program of research on aeroelastic instabilities. The ATW experiment was performed to study a specific instability known as flutter. Flutter is a destructive phenomenon caused by adverse coupling of structural dynamics and aerodynamics. The process of determining a flight envelope within which an aircraft will not experience flutter, known as flight flutter testing, is very dangerous and expensive because predictions of the instability are often unreliable. The ATW was a small-scale airplane wing that comprised an airfoil and boom (see upper part of Figure 1). For flight tests, the ATW was mounted on the F-15B/FTF-II testbed, which is a second-generation flight-test fixture described in Flight-Test Fixture for Aerodynamic Research (DRC- 95-27), NASA Tech Briefs, Vol. 19, No. 9, September 1995, page 84. The ATW was mounted horizontally on this fixture, and the entire assembly was attached to the undercarriage of the F-15B airplane (see lower part of Figure 1). The primary objective of the ATW project was to investigate traditional and advanced methodologies for predicting the onset of flutter. In particular, the ATW generated data that were used to evaluate a flutterometer. This particular flutterometer is an on-line computer program that uses method analysis to estimate worst-case flight conditions associated with flutter. This software was described in A Flutterometer Flight Test Tool NASA Tech Briefs, Vol. 23, No. 1, January 1999, page 52.
The Rake Airflow Gage Experiment was flown on the Propulsion Flight Test Fixture at NASA Dryden Flight Research Center using one of Dryden s F-15B research testbed aircraft. Propulsion Flight Test Fixture is a modular, pylon-based platform for flight testing propulsion system components, such as the Channeled Centerbody Inlet Experiment, an innovative, variable-geometry, mixed compression supersonic inlet under development at NASA Dryden. The objective of this flight test was to ascertain the flowfield angularity and local Mach number profile of the aerodynamic interface plane that is defined by the planned location of the tip of the inlet centerbody. Knowledge of the flowfield characteristics at this location underneath will be essential to computational modeling of the new inlet as well as future propulsion systems flight testing using the test fixture. This paper describes the preparation for and execution of the flight test, as well as results and validation of the algorithm used to calculate local Mach number and angularity from the rake's pressure measurements.
The effects of the Thermal Protection Subsystem (TPS) contamination on the space shuttle orbiter S band quad antenna due to multiple mission buildup are discussed. A test fixture was designed, fabricated and exposed to ten cycles of simulated ground and flight environments. Radiation pattern and impedance tests were performed to measure the effects of the contaminates. The degradation in antenna performance was attributed to the silicone waterproofing in the TPS tiles rather than exposure to the contaminating sources used in the test program. Validation of the accuracy of an analytical thermal model is discussed. Thermal vacuum tests with a test fixture and a representative S band quad antenna were conducted to evaluate the predictions of the analytical thermal model for two orbital heating conditions and entry from each orbit. The results show that the accuracy of predicting the test fixture thermal responses is largely dependent on the ability to define the boundary and ambient conditions. When the test conditions were accurately included in the analytical model, the predictions were in excellent agreement with measurements.
Several potential valve closures for the space shuttle auxiliary propulsion system (SS/APS) were investigated analytically and experimentally in a modeling program. The most promising of these were analyzed and experimentally evaluated in a full-size functional valve test fixture of novel design. The engineering investigations conducted for both model and scale evaluations of the SS/APS valve closures and functional valve fixture are described. Preliminary designs, laboratory tests, and overall valve test fixture designs are presented, and a final recommended flightweight SS/APS valve design is presented.
A flight experiment was conducted to investigate the pressure distribution, local-flow conditions, and boundary-layer transition characteristics on a large flat plate in flight at supersonic speeds up to Mach 2.00. The tests used a NASA testbed aircraft with a bottom centerline mounted test fixture. The primary objective of the test was to characterize the local flow field in preparation for future tests of a high Reynolds number natural laminar flow test article. A second objective was to determine the boundary-layer transition characteristics on the flat plate and the effectiveness of using a simplified surface coating. Boundary-layer transition was captured in both analog and digital formats using an onboard infrared imaging system. Surface pressures were measured on the surface of the flat plate. Flow field measurements near the leading edge of the test fixture revealed the local flow characteristics including downwash, sidewash, and local Mach number. Results also indicated that the simplified surface coating did not provide sufficient insulation from the metallic structure, which likely had a substantial effect on boundary-layer transition compared with that of an adiabatic surface. Cold wall conditions were predominant during the acceleration to maximum Mach number, and warm wall conditions were evident during the subsequent deceleration.
A test fixture for measuring the dynamic performance of candidate high-temperature engine seal concepts was installed at NASA Lewis Research Center. The test fixture was designed to evaluate seal concepts under development for advanced hypersonic engines, such as those being considered for the National Aerospace Plane (NASP). The fixture can measure dynamic seal leakage performance from room temperature up to 840 C (1550 F) and air pressure differentials up to 690 kPa (100 psi). Performance of the seals can be measured while sealing against flat or distorted walls. In the fixture two seals are preloaded against the sides of a 30 cm (1 ft) long saber that slides transverse to the axis of the seals, simulating the scrubbing motion anticipated in these engines. The capabilities of this test fixture along with preliminary data showing the dependence of seal leakage performance on high temperature cycling are addressed.
A lubricant screening test fixture has been devised in order to satisfy the need to obtain lubricant performance data in a timely manner. This fixture has been used to perform short-term tests on potential lubricants for several spacecraft applications. The results of these tests have saved time by producing qualitative performance rankings of lubricant selections prior to life testing. To date, this test fixture has been used to test lubricants for 3 particular applications. The qualitative results from these tests have been verified by life test results and have provided insight into the function of various anti-wear additives.
A test fixture for measuring the dynamic performance of candidate high-temperature engine seal concepts has been installed at NASA Lewis Research Center. The test fixture has been designed to evaluate seal concepts under development for advanced hypersonic engines, such as those being considered for the National Aerospace Plane (NASP). The fixture can measure dynamic seal leakage performance from room temperature up to 840 C (1550 F) and air pressure differentials up to 690 kPa (100 psi). Performance of the seals can be measured while sealing against flat or distorted walls. In the fixture two seals are preloaded against the sides of a 30 cm (1 ft) long saber that slides transverse to the axis of the seals, simulating the scrubbing motion anticipated in these engines. This report covers the capabilities of this test fixture along with preliminary data showing the dependence of seal leakage performance on high temperature cycling.
A parametric study of the effects of test-fixture-induced initial prestress and elastic edge restraints on the prebuckling and buckling responses of a compression-loaded, quasi-isotropic curved panel is presented. The numerical results were obtained by using a geometrically nonlinear finite element analysis code with high-fidelity models. The results presented show that a wide range of prebuckling and buckling behavior can be obtained by varying parameters that represent circumferential loaded-edge restraint and rotational unloaded-edge restraint provided by a test fixture and that represent the mismatch in specimen and test-fixture radii of curvature. For a certain range of parameters, the panels exhibit substantial nonlinear prebuckling deformations that yield buckling loads nearly twice the corresponding buckling load predicted by a traditional linear bifurcation buckling analysis for shallow curved panels. In contrast, the results show another range of parameters exist for which the nonlinear prebuckling deformations either do not exist or are relatively benign, and the panels exhibit buckling loads that are nearly equal to the corresponding linear bifurcation buckling load. Overall, the results should be of particular interest to scientists, engineers, and designers involved in simulating flight-hardware boundary conditions in structural verification and certification tests, involved in validating structural analysis tools, and interested in tailoring buckling performance.
A parametric study of the effects of test-fixture-induced initial prestress and elastic edge restraints on the prebuckling and buckling responses of a compression-loaded, quasi-isotropic curved panel is presented. The numerical results were obtained by using a geometrically nonlinear finite element analysis code with high-fidelity models. The results presented show that a wide range of prebuckling and buckling behavior can be obtained by varying parameters that represent circumferential loaded-edge restraint and rotational unloaded-edge restraint provided by a test fixture and that represent the mismatch in specimen and test-fixture radii of curvature. For a certain range of parameters, the panels exhibit substantial nonlinear prebuckling deformations that yield buckling loads nearly twice the corresponding buckling load predicted by a traditional linear bifurcation buckling analysis for shallow curved panels. In contrast, the results show another range of parameters exist for which the nonlinear prebuckling deformations either do not exist or are relatively benign, and the panels exhibit buckling loads that are nearly equal to the corresponding linear bifurcation buckling load. Overall, the results should also be of particular interest to scientists, engineers, and designers involved in simulating flight-hardware boundary conditions in structural verification and certification tests, involved in validating structural analysis tools, and interested in tailoring buckling performance.