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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.

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At least 847 records · Page 47
All sourcesNASA NTRSDOE OSTI
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.↗
NASA NTRS1992

Environmental qualification testing of the prototype pool boiling experiment

The prototype Pool Boiling Experiment (PBE) flew on the STS-47 mission in September 1992. This report describes the purpose of the experiment and the environmental qualification testing program that was used to prove the integrity of the prototype 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 the project team a wider latitude in determining which shuttle thermal altitudes 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 project's 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 NTRS2000

Peroxide Propulsion at the Turn of the Century

A resurgence of interest in peroxide propulsion has occurred in the last years of the 21st Century. This interest is driven by the need for lower cost propulsion systems and the need for storable reusable propulsion systems to meet future space transportation system architectures. NASA and the Air Force are jointly developing two propulsion systems for flight demonstration early in the 21st Century. One system will be a development of Boeing's AR2-3 engine, which was successfully fielded in the 1960s. The other is a new pressure-fed design by Orbital Sciences Corporation for expendable mission requirements. Concurrently NASA and industry are pursuing the key peroxide technologies needed to design, fabricate, and test advanced peroxide engines to meet the mission needs beyond 2005. This paper will present a description of the AR2-3, report the status of its current test program, and describe its intended flight demonstration. This paper will then describe the Orbital 10K engine, the status of its test program, and describe its planned flight demonstration. Finally the paper will present a plan, or technology roadmap, for the development of an advanced peroxide engine for the 21st Century.

Anderson, William E.↗
NASA NTRS2003

Application of a Split-Fiber Probe to Velocity Measurement in the NASA Research Compressor

A split-fiber probe was used to acquire unsteady data in a research compressor. The probe has two thin films deposited on a quartz cylinder 200 microns in diameter. A split-fiber probe allows simultaneous measurement of velocity magnitude and direction in a plane that is perpendicular to the sensing cylinder, because it has its circumference divided into two independent parts. Local heat transfer considerations indicated that the probe direction characteristic is linear in the range of flow incidence angles of +/- 35. Calibration tests confirmed this assumption. Of course, the velocity characteristic is nonlinear as is typical in thermal anemometry. The probe was used extensively in the NASA Glenn Research Center (GRC) low-speed, multistage axial compressor, and worked reliably during a test program of several months duration. The velocity and direction characteristics of the probe showed only minute changes during the entire test program. An algorithm was developed to decompose the probe signals into velocity magnitude and velocity direction. The averaged unsteady data were compared with data acquired by pneumatic probes. An overall excellent agreement between the averaged data acquired by a split-fiber probe and a pneumatic probe boosts confidence in the reliability of the unsteady content of the split-fiber probe data. To investigate the features of unsteady data, two methods were used: ensemble averaging and frequency analysis. The velocity distribution in a rotor blade passage was retrieved using the ensemble averaging method. Frequencies of excitation forces that may contribute to high cycle fatigue problems were identified by applying a fast Fourier transform to the absolute velocity data.

Lepicovsky, Jan↗
NASA NTRS2008

JWST ISIM Harness Thermal Evaluation

The James Webb Space Telescope (JWST) will be a large infrared telescope with a 6.5-meter primary mirror. Launch is planned for 2013. JWST wl1 be the premier observatory of the next decade serving thousands of astronomers worldwide. The Integrated Science Instrument Module (ISIM) is the unit that will house thc four main JWST instruments. The ISIM enclosure passively cooled to 37 Kelvin and has a tightly managed thermal budget. A significant portion of the ISIM heat load is due to parasitic heat gains from the instrument harnesses. These harnesses provide a thermal path from the Instrument Electronics Control (IEC) to the ISIM. Because of the impact of this load to the ISIM thermal design, understanding the harness parasitic heat gains is critical. To this effect, a thermal test program has been conducted in order to characterize these parasitic loads and verify harness thermal models. Recent parasitic heat loads tests resulted in the addition of a dedicated multiple stage harness radiator. In order for the radiator to efficiently reject heat from the harness, effective thermal contact conductance values for multiple harnesses had to be determined. This presentation will describe the details and the results of this test program.

Kobel, Mark↗
NASA NTRS1982

Full-scale flammability test data for validation of aircraft fire mathematical models

Twenty-five large scale aircraft flammability tests were conducted in a Boeing 737 fuselage at the NASA Johnson Space Center (JSC). The objective of this test program was to provide a data base on the propagation of large scale aircraft fires to support the validation of aircraft fire mathematical models. Variables in the test program included cabin volume, amount of fuel, fuel pan area, fire location, airflow rate, and cabin materials. A number of tests were conducted with jet A-1 fuel only, while others were conducted with various Boeing 747 type cabin materials. These included urethane foam seats, passenger service units, stowage bins, and wall and ceiling panels. Two tests were also included using special urethane foam and polyimide foam seats. Tests were conducted with each cabin material individually, with various combinations of these materials, and finally, with all materials in the cabin. The data include information obtained from approximately 160 locations inside the fuselage.

Kuminecz, J. F.↗
NASA NTRS1985

Critical joints in large composite primary aircraft structures. Volume 1: Technical summary

A program was conducted at Douglas Aircraft Company to develop the technology for critical joints in composite wing structure that meets all the design requirements of a 1990 commercial transport aircraft. In fulfilling this objective, analytical procedures for joint design and analysis were developed during Phase 1 of the program. Tests were conducted at the element level to supply the empirical data required for methods development. Large composite multirow joints were tested to verify the selected design concepts and for correlation with analysis predictions. The Phase 2 program included additional tests to provide joint design and analysis data, and culminated with several technology demonstration tests of a major joint area representative of a commercial transport wing. The technology demonstration program of Phase 2 is discussed. The analysis methodology development, structural test program, and correlation between test results and analytical strength predictions are reviewed.

Bunin, Bruce L.↗
NASA NTRS1981

American Airlines LOFT evaluation program

The development of a test program to evaluate recurrent training LOFT and a three-legged scenario used for the evaluation are highlighted. The test guidelines set up and the questionnaires sent to crew member participants are examined.

Jensen, D.↗
NASA NTRS1988

Study of toluene rotary fluid management device and shear flow condenser performance for a space-based organic Rankine power system

Management of two-phase fluid and control of the heat transfer process in microgravity is a technical challenge that must be addressed for an orbital Organic Rankine Cycle (ORC) application. A test program was performed in 1-g that satisfactorily demonstrated the two-phase management capability of the rotating fluid management device (RFMD) and shear-flow condenser. Operational tests of the RFMD and shear flow condenser in adverse gravity orientations, confirmed that the centrifugal forces in the RFMD and the shear forces in the condenser were capable of overcoming gravity forces. In a microgravity environment, these same forces would not have to compete against gravity and would therefore be dominant. The specific test program covered the required operating range of the Space Station Solar Dynamic Rankine Cycle power system. Review of the test data verified that: fluid was pumped from the RFMD in all attitudes; subcooled states in the condenser were achieved; condensate was pushed uphill against gravity; and noncondensible gases were swept through the condenser.

Havens, Vance↗
NASA NTRS1991

Hypersonic Research Vehicle (HRV) real-time flight test support feasibility and requirements study. Part 2: Remote computation support for flight systems functions

The requirements are assessed for the use of remote computation to support HRV flight testing. First, remote computational requirements were developed to support functions that will eventually be performed onboard operational vehicles of this type. These functions which either cannot be performed onboard in the time frame of initial HRV flight test programs because the technology of airborne computers will not be sufficiently advanced to support the computational loads required, or it is not desirable to perform the functions onboard in the flight test program for other reasons. Second, remote computational support either required or highly desirable to conduct flight testing itself was addressed. The use is proposed of an Automated Flight Management System which is described in conceptual detail. Third, autonomous operations is discussed and finally, unmanned operations.

Rediess, Herman A.↗