Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “TEST PROGRAM”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 865 records · Page 48
All sourcesNASA NTRSDOE OSTI
NASA NTRS2000

On Fast Post-Processing of Global Positioning System Simulator Truth Data and Receiver Measurements and Solutions Data

Post-processing of data, related to a GPS receiver test in a GPS simulator and test facility, is an important step towards qualifying a receiver for space flight. Although the GPS simulator provides all the parameters needed to analyze a simulation, as well as excellent analysis tools on the simulator workstation, post-processing is not a GPS simulator or receiver function alone, and it must be planned as a separate pre-flight test program requirement. A GPS simulator is a critical resource, and it is desirable to move off the pertinent test data from the simulator as soon as a test is completed. The receiver and simulator databases are used to extract the test data files for postprocessing. These files are then usually moved from the simulator and receiver systems to a personal computer (PC) platform, where post-processing is done typically using PC-based commercial software languages and tools. Because of commercial software systems generality their functions are notoriously slow and more than often are the bottleneck even for short duration simulator-based tests. There is a need to do post-processing faster and within an hour after test completion, including all required operations on the simulator and receiver to prepare and move off the post-processing files. This is especially significant in order to use the previous test feedback for the next simulation setup or to run near back-to-back simulation scenarios. Solving the post-processing timing problem is critical for a pre-flight test program success. Towards this goal an approach was developed that allows to speed-up post-processing by an order of a magnitude. It is based on improving the post-processing bottleneck function algorithm using a priory information that is specific to a GPS simulation application and using only the necessary volume of truth data. The presented postprocessing scheme was used in support of a few successful space flight missions carrying GPS receivers.

Kizhner, Semion↗
NASA NTRS2003

Implementation of Wireless and Intelligent Sensor Technologies in the Propulsion Test Environment

From the first Saturn V rocket booster (S-II-T) testing in 1966 and the routine Space Shuttle Main Engine (SSME) testing beginning in 1975, to more recent test programs such as the X-33 Aerospike Engine, the Integrated Powerhead Development (IPD) program, and the Hybrid Sounding Rocket (HYSR), Stennis Space Center (SSC) continues to be a premier location for conducting large-scale propulsion testing. Central to each test program is the capability for sensor systems to deliver reliable measurements and high quality data, while also providing a means to monitor the test stand area to the highest degree of safety and sustainability. As part of an on-going effort to enhance the testing capabilities of Stennis Space Center, the Test Technology and Development group is developing and applying a number of wireless and intelligent sensor technologies in ways that are new to the test existing test environment.

Solano, Wanda M.↗
NASA NTRS1993

Flight-determined engine exhaust characteristics of an F404 engine in an F-18 airplane

Personnel at the NASA Langley Research Center (NASA-Langley) and the NASA Dryden Flight Research Facility (NASA-Dryden) recently completed a joint acoustic flight test program. Several types of aircraft with high nozzle pressure ratio engines were flown to satisfy a twofold objective. First, assessments were made of subsonic climb-to-cruise noise from flights conducted at varying altitudes in a Mach 0.30 to 0.90 range. Second, using data from flights conducted at constant altitude in a Mach 0.30 to 0.95 range, engineers obtained a high quality noise database. This database was desired to validate the Aircraft Noise Prediction Program and other system noise prediction codes. NASA-Dryden personnel analyzed the engine data from several aircraft that were flown in the test program to determine the exhaust characteristics. The analysis of the exhaust characteristics from the F-18 aircraft are reported. An overview of the flight test planning, instrumentation, test procedures, data analysis, engine modeling codes, and results are presented.

Ennix, Kimberly A.↗
NASA NTRS1973

An investigation of a sterile access technique for the repair and adjustment of sterile spacecraft

A description is presented of a unique system for the sterilization and sterile repair of spacecraft and the results of a test program designed to assess the biological integrity and engineering reliability of the system. This trailer-mounted system, designated the model assembly sterilizer for testing (MAST), is capable of the dry-heat sterilization of spacecraft and/or components less than 2.3 meters in diameter at temperatures up to 433 K and the steam sterilization of components less than 0.724 meter in diameter. Sterile access to spacecraft is provided by two tunnel suits, called the bioisolator suit systems (BISS), which are contiguous with the walls of the sterilization chambers. The test program was designed primarily to verify the biological and engineering reliability of the MAST system by processing simulated space hardware. Each test cycle simulated the initial sterilization of a spacecraft, sterile repair of a failed component, removal of the spacecraft from the MAST for mating with the bus, and a sterile recycle repair.

Farmer, F. H.↗
NASA NTRS2008

A Capable and Temporary Test Facility on a Shoestring Budget: The MSL Touchdown Test Facility

The Mars Science Laboratory mission (MSL) has undertaken a developmental Touchdown Test Program that utilizes a full-scale rover vehicle and an overhead winch system to replicate the skycrane landing event. Landing surfaces consisting of flat and sloped granular media, planar, rigid surfaces, and various combinations of rocks and slopes were studied. Information gathered from these tests was vital for validating the rover analytical model, validating certain design or system behavior assumptions, and for exploring events and phenomenon that are either very difficult or too costly to model in a credible way. This paper describes this test program, with a focus on the creation of test facility, daily test operations, and some of the challenges faced and lessons learned along the way.

rover↗
NASA NTRS1978

Space Shuttle Orbiter Approach and Landing Test: Final Evaluation Report

The Approach and Landing Test Program consisted of a series of steps leading to the demonstration of the capability of the Space Shuttle orbiter to safely approach and land under conditions similar to those planned for the final phases of an orbital flight. The tests were conducted with the orbiter mounted on top of a specially modified carrier aircraft. The first step provided airworthiness and performance verification of the carrier aircraft after modification. The second step consisted of three taxi tests and five flight tests with an inert unmanned orbiter. The third step consisted of three mated tests with an active manned orbiter. The fourth step consisted of five flights in which the orbiter was separated from the carrier aircraft. For the final two flights, the orbiter tail cone was replaced by dummy engines to simulate the actual orbital configuration. Landing gear braking and steering tests were accomplished during rollouts following the free flight landings. Ferry testing was integrated into the Approach and Landing Test Program to the extent possible. In addition, four ferry test flights were conducted with the orbiter mated to the carrier aircraft in the ferry configuration after the free-flight tests were completed.

Source record↗
NASA NTRS2013

International Space Station (ISS) Soyuz Vehicle Descent Module Evaluation of Thermal Protection System (TPS) Penetration Characteristics

The descent module (DM) of the ISS Soyuz vehicle is covered by thermal protection system (TPS) materials that provide protection from heating conditions experienced during reentry. Damage and penetration of these materials by micrometeoroid and orbital debris (MMOD) impacts could result in loss of vehicle during return phases of the mission. The descent module heat shield has relatively thick TPS and is protected by the instrument-service module. The TPS materials on the conical sides of the descent module (referred to as backshell in this test plan) are exposed to more MMOD impacts and are relatively thin compared to the heat shield. This test program provides hypervelocity impact (HVI) data on materials similar in composition and density to the Soyuz TPS on the backshell of the vehicle. Data from this test program was used to update ballistic limit equations used in Soyuz TPS penetration risk assessments. The impact testing was coordinated by the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) Group [1] in Houston, Texas. The HVI testing was conducted at the NASA-JSC White Sands Hypervelocity Impact Test Facility (WSTF) at Las Cruces, New Mexico. Figure

Davis, Bruce A.↗
NASA NTRS1992

Design and analysis of seals for extended service life

Space Station Freedom is being developed for a service life of up to thirty years. As a consequence, the design requirements for the seals to be used are unprecedented. Full scale testing to assure the selected seals can satisfy the design requirements are not feasible. As an alternative, a sub-scale test program has been developed by MSFC to calibrate the analysis tools to be used to certify the proposed design. This research has been conducted in support of the MSFC Integrated Seal Test Program. The ultimate objective of this research is to correlate analysis and test results to qualify the analytical tools, which in turn, are to be used to qualify the flight hardware. This research is totally focused on O-rings that are compressed by perpendicular clamping forces. In this type of seal the O-ring is clamped between the sealing surfaces by loads perpendicular to the circular cross section.

Bower, Mark V.↗
NASA NTRS1992

Environmental qualification testing of payload G-534, the Pool Boiling Experiment

Payload G-534, the prototype Pool Boiling Experiment (PBE), is scheduled to fly on the STS-47 mission in September 1992. This paper describes the purpose of the experiment and the environmental qualification testing program that was used to prove the integrity of the hardware. Component and box level vibration and thermal cycling tests were performed to give an early level of confidence in the hardware designs. At the system level, vibration, thermal extreme soaks, and thermal vacuum cycling tests were performed to qualify the complete design for the expected shuttle environment. The system level vibration testing included three axis sine sweeps and random inputs. The system level hot and cold soak tests demonstrated the hardware's capability to operate over a wide range of temperatures and gave wider latitude in determining which shuttle thermal attitudes were compatible with the experiment. The system level thermal vacuum cycling tests demonstrated the hardware's capability to operate in a convection free environment. A unique environmental chamber was designed and fabricated by the PBE team and allowed most of the environmental testing to be performed within the hardware build laboratory. The completion of the test program gave the project team high confidence in the hardware's ability to function as designed during flight.

Sexton, J. Andrew↗
NASA NTRS1997

Mars Pathfinder Spacecraft, Lander, and Rover Testing in Simulated Deep Space and Mars Surface Environments

The Mars Pathfinder (MPF) Spacecraft was built and tested at the Jet Propulsion Laboratory during 1995/96. MPF is scheduled to launch in December 1996 and to land on Mars on July 4, 1997. The testing program for MPF required subjecting the mission hardware to both deep space and Mars surface conditions. A series of tests were devised and conducted from 1/95 to 7/96 to study the thermal response of the MPF spacecraft to the environmental conditions in which it will be exposed during the cruise phase (on the way to Mars) and the lander phase (landed on Mars) of the mission. Also, several tests were conducted to study the thermal characteristics of the Mars rover, Sojourner, under Mars surface environmental conditions. For these tests, several special test fixtures and methods were devised to simulate the required environmental conditions. Creating simulated Mars surface conditions was a challenging undertaking since Mars' surface is subjected to diurnal cycling between -20 C and -85 C, with windspeeds to 20 m/sec, occurring in an 8 torr CO2 atmosphere. This paper describes the MPF test program which was conducted at JPL to verify the MPF thermal design.

Johnson, Kenneth R.↗
NASA NTRS2008

Preliminary Analysis of Acoustic Measurements from the NASA-Gulfstream Airframe Noise Flight Test

The NASA-Gulfstream joint Airframe Noise Flight Test program was conducted at the NASA Wallops Flight Facility during October, 2006. The primary objective of the AFN flight test was to acquire baseline airframe noise data on a regional jet class of transport in order to determine noise source strengths and distributions for model validation. To accomplish this task, two measuring systems were used: a ground-based microphone array and individual microphones. Acoustic data for a Gulfstream G550 aircraft were acquired over the course of ten days. Over twenty-four test conditions were flown. The test matrix was designed to provide an acoustic characterization of both the full aircraft and individual airframe components and included cruise to landing configurations. Noise sources were isolated by selectively deploying individual components (flaps, main landing gear, nose gear, spoilers, etc.) and altering the airspeed, glide path, and engine settings. The AFN flight test program confirmed that the airframe is a major contributor to the noise from regional jets during landing operations. Sound pressure levels from the individual microphones on the ground revealed the flap system to be the dominant airframe noise source for the G550 aircraft. The corresponding array beamform maps showed that most of the radiated sound from the flaps originates from the side edges. Using velocity to the sixth power and Strouhal scaling of the sound pressure spectra obtained at different speeds failed to collapse the data into a single spectrum. The best data collapse was obtained when the frequencies were left unscaled.

Khorrami, Mehdi R.↗
NASA NTRS1993

Viscoelastic analysis of seals for extended service life

The space station is being developed for a service life of up to thirty years. As a consequence, the design requirements for the seals to be used are unprecedented. Full scale testing to assure the selected seals can satisfy the design requirements are not feasible. As an alternative, a sub-scale test program (2) has been developed by MSFC to calibrate the analysis tools to be used to certify the proposed design. This research has been conducted in support of the MSFC Integrated Seal Test Program. The ultimate objective of this research is to correlate analysis and test results to qualify the analytical tools which in turn, are to be used to qualify the flight hardware. Seals are simple devices, in wide spread use. The most common type of seal is the O-ring. O-ring seals are typically rings of rubber with a circular cross section. The rings are placed between the surfaces to be sealed, usually in a groove of some design. The particular design may differ based on a number of different factors. This research is focused on O-rings that are staticly compressed by perpendicular clamping forces, commonly referred to as face seals. In this type of seal the O-ring is clamped between the sealing surfaces by loads perpendicular to the circular cross section.

Bower, Mark V.↗
NASA NTRS1964

Thermal testing of the Ranger Block III SPACECRAFT in the JPL 25 ft. space simulator

In January, 1964, a test program was begun on the thermal design of the Ranger Block III Spacecraft. The tests were performed i n the newly operational JPL 25' Space Simulator over a period of 6 months . The objectives of these tests were two-fold: A. To evaluate the 25' Space Simulator as a facility for proving the thermal design of spacecrafts, and B. To verify .the thermal design of the Ranger Block III spacecraft. These two objectives are complimentary in the test series performed and are difficult to separate into distinct categories. An important part of the first objective was to learn what type of test preparation, instrumentation, and analysis was required for the meaningful evaluation of test data from the 25' Space Simulator tests. Although some experience had been gained in testing of components and incomplete spacecrafts in smaller solar simulation chambers during the early part of the Ranger program, we knew little about testing of a complete spacecraft in the 25' Space Simulator when this test program began. Test analysis requires that the energy absorbed by various spacecraft components be known. This requires a knowledge of the area of solar absorption or sunlit area, the solar energy flux density on the area, and the effective absorptance of that area. Most of the problems encountered were associated with the determination of the last two quantities since the sunlit area may be obtained directly by inspection of spacecraft surfaces. This paper will present a discussion of our experiences during the Ranger Block III thermal test series on the Thermal Test Model (TCM). The TCM was thermally equivalent to the flight type Ranger spacecraft except for the lack of an antenna dish and solar panels. Flight type structural hardware was used with surface finishes equivalent to those of the flight spacecraft. Aluminum blocks simulated the spacecraft electronics thermal masses with resistance heaters simulating the electronic power dissipation. The discussion will be presented in a semi-chronological order and will be divided into the following areas: 1. Determination of solar simulation flux density on spacecraft surfaces. 2. Problems related to decollimation of the solar simulation source. 3. Determination of effective absorptance in the solar simulation spectrum.

SPACE SIMULATOR↗
NASA NTRS1990

Hot gas ingestion characteristics and flow visualization of a vectored thrust STOVL concept

A 9.2 percent scale short takeoff and vertical landing (STOVL) hot gas ingestion model was designed and built by McDonnell Douglas Corporation (MCAIR) and tested in the NASA Lewis Research Center 9- by 15-Foot Low Speed Wind Tunnel (LSWT). Hot gas ingestion, the entrainment of heated engine exhaust into the inlet flow field, is a key development issue for advanced short takeoff and vertical landing aircraft. The Phase 1 test program, conducted by NASA Lewis and McDonnell Douglas Corporation, evaluated the hot ingestion phenomena and control techniques and Phase 2 test program which was conducted by NASA Lewis are both reported. The Phase 2 program was conducted at exhaust nozzles temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/lift improvement devices which reduced the hot gas ingestion. The model support system had four degrees of freedom, heated high pressure air for nozzle flow, and a suction system exhaust for inlet flow. The headwind (freestream) velocity for Phase 1 was varied from 8 to 90 kn, with primary data taken in the 8 to 23 kn headwind velocity range. Phase 2 headwind velocity varied from 10 to 23 kn. Results of both Phase 1 and 2 are presented. A description of the model, facility, a new model support system, and a sheet laser illumination system are also provided. Results are presented over a range of main landing gear height (model height) above the ground plane at a 10 kn headwind velocity. The results contain the compressor face pressure and temperature distortions, total pressure recovery, compressor face temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane temperature and pressure distributions, model airframe heating, and the location of the ground flow separation. Results from the sheet laser flow visualization test are also shown.

Johns, Albert L.↗
NASA NTRS1993

Development of radiometer windows for atmospheric entry vehicles

A test program was established to determine the feasibility of developing a radiometer window for stagnation region shock layer atmospheric entry experiments from geosynchronous orbit and with measurements extending into the VUV region to about 170 nm. This test program examined both isolated (uncooled) and cooled window mounting options. Preliminary tests determined the gross suitability of sapphire and fused silica as window materials. Measurements show that the expected maximum temperature (1500 C) is itself not a problem, but that temperature gradients can easily break a window and that the transmission of sapphire and of fused silica at 174 nm drops dramatically at temperatures beyond 800 C and 500 C, respectively. A series of developmental tests showed that the surface of an uncooled fused silica window begins to melt at the entry conditions tested while an uncooled sapphire window generally does not, and that a cooled sapphire window design that keeps the window below 800 C appears possible. The final proof-of-design tests narrowed the selection of possible candidate window sizes and mounting configurations. These tests demonstrate that the best insulated and cooled window designs use small diameter sapphire windows and that thermal modeling codes can be used to assist the design process.

Terrazas-Salinas, Imelda↗
NASA NTRS2010

Performance and Safety Tests of Lithium-Ion Cells Arranged in a Matrix Design Configuration

Matrix Packs display large variations in cell bank voltages at the charge and discharge current (C/2) used in this test program. The voltage difference is larger at the end of discharge than at the end of charge under the conditions studied. Disconnection of a cell from the pack leads to a larger voltage difference during discharge (greater than 2.0 V) between the bank that has one less cell and the other banks. Thermal profile does not show any significant changes or increase in temperature after one cell was disconnected from the bank in spite of falling to very low voltages at the end of discharge. All tests on the matrix pack with the HAM displayed lower max in general due to the placement of thermocouple on the outside of the HAM rather than on the cells. Disconnection of cells has almost no influence on the performance of the packs and does not show any abnormal thermal changes for the 100 cycles obtained in this test program. Longer cycle life may influence the performance especially if the low voltage cell goes into reversal. Overcharge leads to CID activation of cells. If the matrix configuration has a larger number of cells in series, (more than 5 S configuration), the limitations of protective devices may manifest itself irrespective of it being in a matrix configuration. External short circuit causes a fire with expulsion of content from some cells. The fire does not propagate itself laterally, but if there was cell module stacking, then the fire would cause the cells above it to also go into flames/thermal runaway. Limitations of protective devices are observed in this case as the PTCs in the cells did not protect under this abusive condition. Matrix configurations seem to provide protection against lateral propagation of fire and flame. Matrix pack configuration seems to provide good performance in spite of losing cell connections; at least for the configuration tested under this program.

Jeevarajan, Judith↗
NASA NTRS1977

The remote measurement of tornado-like flows employing a scanning laser Doppler system

The paper deals with a scanning laser Doppler velocimeter system employed in a test program for measuring naturally occurring tornado-like phenomena, known as dust devils. A description of the system and the test program is followed by a discussion of the data processing techniques and data analysis. The system uses a stable 15-W CO2 laser with the beam expanded and focused by a 12-inch telescope. Range resolution is obtained by focusing the optical system. The velocity of each volume of air (scanned in a horizontal plane) is determined from spectral analysis of the heterodyne signal. Results derived from the measurement program and data/system analyses are examined.

Jeffreys, H. B.↗
NASA NTRS1985

The space shuttle launch vehicle aerodynamic verification challenges

The Space Shuttle aerodynamics and performance communities were challenged to verify the Space Shuttle vehicle (SSV) aerodynamics and system performance by flight measurements. Historically, launch vehicle flight test programs which faced these same challenges were unmanned instrumented flights of simple aerodynamically shaped vehicles. However, the manned SSV flight test program made these challenges more complex because of the unique aerodynamic configuration powered by the first man-rated solid rocket boosters (SRB). The analyses of flight data did not verify the aerodynamics or performance preflight predictions of the first flight of the Space Transportation System (STS-1). However, these analyses have defined the SSV aerodynamics and verified system performance. The aerodynamics community also was challenged to understand the discrepancy between the wind tunnel and flight defined aerodynamics. The preflight analysis challenges, the aerodynamic extraction challenges, and the postflight analyses challenges which led to the SSV system performance verification and which will lead to the verification of the operational ascent aerodynamics data base are presented.

Wallace, R. O.↗