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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 1,603 records · Page 89
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

ROLM computers in the flight testing of the Fokker F29 aircraft

The design, development, installation, and operation of the test equipment (MRVS) is discussed. The main feature of the MRVS is continuous recording on an instrumentation recorder data of up to 1500 parameters with a total sample rate of up to 10,000 samples per second. In order to compress the evaluation and calibration time period, the following additional requirements were set for two test systems: (1) recording of selected parameters, time tagged on computer compatible tape (CCT); (2) recording of selected high bandwidth signals and ad hoc parameters on analog tape; (3) onboard presentation of calibrated parameter data, in engineering units, in numerical as well as graphical form for: system checkout during pre-, in- and post-flight, and quick look analysis during in-flight; (4) real time presentation on the ground by telemetry for: flight monitoring, takeoff and landing measurements, and noise measurements; and (5) data processing on the ground on the Fokker-NLR computer network.

Manders, P. J.↗
NASA NTRS1995

Spacecraft Availability Enhancement by In-flight Testing of Spare Parts

This paper describes an approach to improve availability by testing redundant parts at pre-determined intervals. The purpose of the testing is to detect nonfunctional back-up equipment and develop work-around measures or replacement spacecraft before a failure of the primary equipment reduces availability. The work reported here is an outgrowth of the NASA Space Network's program for the maintenance and replenishment of the Tracking and Data Relay Satellite System. The approach is based on a standby factor and a cyclic stress factor. The standby factor accounts for the effects of adverse storage conditions encountered as part of the in-flight environment. The stress factor accounts for the effects of physical or thermal cycling due to the application of force or power that characterizes the operation or use of a component. By the quantitative consideration of standby and cycling risks, a regular testing interval can be calculated - an interval for testing that furnishes information for availability planning but does not subject the spacecraft to undue risk. This paper includes quantitative solutions for appropriate testing intervals for equipment configurations and failure rates that are representative of a Tracking and Relay Data Satellite. The effect of a single test on the availability of certain equipment is also illustrated.

Cutler, Robert M.↗
NASA NTRS1995

Final Results of the Advanced Photovoltaic Experiment Flight Test

The Advanced Photovoltaic Experiment was designed to generate laboratory reference standards as well as to explore the durability of a wide variety of space solar cells. In addition to the cells, it was equipped with an absolute cavity radiometer to measure solar intensity, a spectroradiometer to measure the spectral content of this radiation and a sun angle sensor. Data from the solar cells and various sensors was obtained on a daily basis during the first eleven months of the 69 month flight. In this paper pre-flight and post-flight laboratory measurements are compared with on-orbit calibration data. Pre-flight and post-flight calibration data of the cavity radiometer as well as on-orbit data demonstrated the accuracy and durability of the Eppley instrument flown on APEX.

Brinker, David J.↗
NASA NTRS2012

Overheating Anomalies during Flight Test Due to the Base Bleeding

In this paper we present the results of the analytical and numerical studies of the plume interaction with the base flow in the presence of base out-gassing. The physics-based analysis and CFD modeling of the base heating for single solid rocket motor performed in this research addressed the following questions: what are the key factors making base flow so different from that in the Shuttle [1]; why CFD analysis of this problem reveals small plume recirculation; what major factors influence base temperature; and why overheating was initiated at a given time in the flight. To answer these questions topological analysis of the base flow was performed and Korst theory was used to estimate relative contributions of radiation, plume recirculation, and chemically reactive out-gassing to the base heating. It was shown that base bleeding and small base volume are the key factors contributing to the overheating, while plume recirculation is effectively suppressed by asymmetric configuration of the flow formed earlier in the flight. These findings are further verified using CFD simulations that include multi-species gas environment both in the plume and in the base. Solid particles in the exhaust plume (Al2O3) and char particles in the base bleeding were also included into the simulations and their relative contributions into the base temperature rise were estimated. The results of simulations are in good agreement with the temperature and pressure in the base measured during the test.

base bleeding↗
NASA NTRS2018

Orion Heat Shield Manufacturing Producibility Improvements for the EM-1 Flight Test Program

This paper describes how the ORION program is incorporating improvements in the heat shield design and manufacturing processes reducing programmatic risk and ensuring crew safety in support of NASA's Exploration missions. The approach for the EFT-1 heat shield utilized a low risk Apollo heritage design and manufacturing process using an Avcoat TPS ablator with a honeycomb substrate to provide a one piece heat shield to meet the mission re-entry heating environments. The EM-1 mission will have additional flight systems installed to fly to the moon and return to Earth. Heat shield design and producibility improvements have been incorporated in the EM-1 vehicle to meet deep space mission requirements. The design continues to use the Avcoat material, but in a block configuration to enable improvements in consistant and repeatable application processes using tile bonding experience developed on the Space Shuttle Transportation System Program.

Producibility↗
DOE OSTI2023

Test Flights of NOAA Ozone Photometer for Uncrewed Aerial Systems (UAS)/Balloon-Based Studies (OPUS) during TRACER (Field Campaign Report)

We proposed to integrate and fly the National Oceanic and Atmospheric Administration (NOAA) Ozone Photometer for Uncrewed Aerial Systems (UAS)/Balloon-Based Studies (OPUS) on the U.S. Department of Energy (DOE) Atmospheric Radiation Measurement (ARM) user facility’s tethered balloon system (TBS) during the second phase (August 1-15) of the 2022 Tracking Aerosol Convection Interactions Experiment (TRACER) in southwest Houston, Texas. The two main goals were: 1) Test fly the sensor to discover and address any potential problems such as interface and communication with other instruments, interferences with other instruments and radio, and 2) If OPUS performs well, then inter-compare OPUS with the EN-SCI ozonesondes that will also be flown on the TBS. The collaborating institutions are DOE Sandia National Laboratories TBS team, the University of Houston, and St. Edward's University.

47 OTHER INSTRUMENTATION↗
NASA NTRS1974

Reynolds number effects on boattail drag of exhaust nozzles from wind tunnel and flight tests

A family of nacelle mounted high angle boattail nozzles was tested to investigate Reynolds number effects on drag. The nozzles were flown on a modified F-106B and mounted on scale models of a F-106 in a wind tunnel. A 19- to 1-range of Reynolds number was covered as a result of the large size differences between models and by flying over a range of altitude. In flight, the nozzles were mounted behind J-85 turbojet engines. Jet boundary simulators and a powered turbojet engine simulator were used on the wind tunnel models. Data were taken at Mach numbers of 0.6 and 0.9. Boattail drag was found to be affected by boattail number. The effect is a complex relationship dependent upon boundary layer thickness and nozzle boattail shape. As Reynolds number was increased from the lowest values obtained with scale models, boattail drag first increased to a maximum at the lowest flight Reynolds number and then decreased.

Wilcox, F. A.↗
NASA NTRS1974

Reynolds number effects on boattail drag of exhaust nozzles from wind tunnel and flight tests

A family of nacelle-mounted high-angle boattail nozzles was tested to investigate Reynolds number effects on drag. The nozzles were flown on a modified F-106B and mounted on scale models of an F-106 in a wind tunnel. A 19-to-1 range of Reynolds number was covered as a result of the large size differences between models and by flying over a range of altitude. In flight the nozzles were mounted behind J-85 turbojet engines. Jet boundary simulators and a powered turbojet engine simulator were used on the wind tunnel models. Data were taken at Mach numbers of 0.6 and 0.9. Boattail drag was found to be affected by Reynolds number. The effect is a complex relationship dependent upon boundary layer thickness and nozzle boattail shape. As Reynolds number was increased from the lowest values obtained with scale models, boattail drag first increased to a maximum at the lowest flight Reynolds number and then decreased.

Wilcox, F. A.↗
NASA NTRS1975

Reynolds number effects on boattail drag of exhaust nozzles from wind tunnel and flight tests

A family of nacelle mounted high angle boattail nozzles was tested to investigate Reynolds number effects on drag. The nozzles were flown on a modified F-106B and mounted on scale models of an F-106 in a wind tunnel. A 19- to 1-range of Reynolds number was covered as a result of the large size differences between models and by flying over a range of altitude. In flight the nozzles were mounted behind J-85 turbojet engines. Jet boundary simulators and a powered turbojet engine simulator were used on the wind tunnel models. Data were taken at Mach numbers of 0.6 and 0.9. Boattail drag was found to be affected by Reynolds number. The effect is a complex relationship dependent upon boundary layer thickness and nozzle boattail shape. As Reynolds number was increased from the lowest values obtained with scale models, boattail drag first increased to a maximum at the lowest flight Reynolds number and then decreased.

Wilcox, F. A.↗
NASA NTRS1975

Analysis of wake vortex flight test data behind a T-33 aircraft

Measurements of the vortex system behind a T-33 aircraft were obtained by a Learjet equipped with a boom carrying a three-wire, hot-wire anemometry probe and other instrumentation. Analysis of the measurements using a computerized geometric method indicated the vortices had a core radius of approximately 0.11 meter with a maximum velocity of 25 meters per second. The hot-wire anemometer was found to be a practical and sensitive instrument for determining in-flight vortex velocities. No longitudinal instabilities, buoyant effects or vortex breakdowns were evident in the data which included vortex wake cross sections from 0.24 to 5.22 kilometers behind the T-33.

Kuhn, G. D.↗
NASA NTRS1978

Ride quality flight testing

The ride quality experienced by passengers is a function of airframe rigid-body, elastic dynamic responses, autopilot, and stability augmentation system control inputs. A frequency response method has been developed to select sinusoidal elevator input time histories yielding vertical load factor distributions, within a given limit, as a function of fuselage station. The numerical technique is illustrated by applying two-degree-of-freedom short-period and first symmetric mode equations of motion to a B-1 aircraft at Mach 0.85 during sea level flight conditions.

Swaim, R. L.↗
NASA NTRS1981

Jet noise results from static, wind tunnel, and flight tests of conical and mechanical suppressor nozzles

Results of jet noise suppression tests conducted on a Rolls-Royce Viper 601 turbojet engine are reported. Seven exhaust nozzle configurations are tested, including two conical nozzles, two suppressor mixers, and three treated ejector configurations with different ejector inlets. Tests are conducted at the NASA Ames outdoor static test facility and the 40- by 80-ft wind tunnel facility at minimum tunnel flow velocity and normal flow velocities of 230 and 290 ft/sec. Near-field multiple sideline noise levels are projected to the far fields to compare far-field fixed microphone outdoor static noise levels, and wind tunnel near-field noise data are projected to the far field and flight distances to compare with noise levels recorded from an Hs-125 aircraft. Near-field outdoor noise data duplicate the far-field data recorded from fixed microphones within 2 PNdB, and the Douglas mechanical jet noise suppressor/treated ejector exhaust system achieves a noise reduction of 12 EPNdB relative to a conic reference nozzle at equal thrust in flight.

Mckinnon, R. A.↗
NASA NTRS2019

Radar Detection of High Concentrations of Ice Particles - Methodology and Preliminary Flight Test Results

High Ice Water Content (HIWC) has been identified as a primary causal factor in numerous engine events over the past two decades. Previous attempts to develop a remote detection process utilizing modern commercial radars have failed to produce reliable results. This paper discusses the reasons for previous failures and describes a new technique that has shown very encouraging accuracy and range performance without the need for any hardware modifications to industry’s current radar designs. The performance of this new process was evaluated during the joint NASA/FAA HIWC RADAR II Flight Campaign in August of 2018. Results from that evaluation are discussed, along with the potential for commercial application, and development of minimum operational performance standards for a future commercial radar product.

Harrah, Steven D.↗
NASA NTRS1980

Analysis of flight test measurements in ground effect

Three sources of errors, introduced into the air-data measurements of altitude and airspeed as a result of ground proximity, are evaluated. It is shown that the primary error results from ground constraint of the wing lifting pattern. Smaller errors result from constraint of the flow over the fuselage and of the engine exhaust. Equations are derived to provide corrections for the three error contributions. The equations are general and can be applied to airplanes of different geometries operating in different conditions. Applications illustrating different types of flight operations (takeoff and landing) are presented using CV-990 data. Applications illustrating the effect of different static orifice locations (fuselage and nose-boom) are presented with YC-15 data. Finally, an application to trajectory reconstruction is illustrated using DC-10 data.

Parks, E. K.↗
NASA NTRS1994

High-angle-of-attack yawing moment asymmetry of the X-31 aircraft from flight test

Significant yawing moment asymmetries were encountered during the high-angle-of-attack envelope expansion of the two X-31 aircraft. These asymmetries led to position saturations of the thrust vector vanes and trailing-edge flaps during some of the dynamic stability axis rolling maneuvers at high angles of attack. This slowed the high-angle-of-attack envelope expansion and resulted in maneuver restrictions. Several aerodynamic modifications were made to the X-31 forebody with the goal of minimizing the asymmetry. A method for determining the yawing moment asymmetry from flight data was developed and an analysis of the various configuration changes completed. The baseline aircraft were found to have significant asymmetries above 45 deg angle of attack with the largest asymmetry typically occurring around 60 deg angle of attack. Applying symmetrical boundary layer transition strips along the forebody sides increased the magnitude of the asymmetry and widened the angle-of-attack range over which the largest asymmetry acted. Installing longitudinal forebody strakes and rounding the sharp nose of the aircraft caused the yawing moment asymmetry magnitude to be reduced. The transition strips and strakes made the asymmetry characteristic of the aircraft more repeatable than the clean forebody configuration. Although no geometric differences between the aircraft were known, ship 2 consistently had larger yawing moment asymmetries than ship 1.

Cobleigh, Brent R.↗
NASA NTRS1962

Free-Flight Tests of Fifth-Stage Scout Entry Vehicle at Mach Numbers of 5 and 17

Measurements have been made in air at two Mach numbers of the static stability, normal force, and drag of a version of the fifth-stage Scout entry vehicle. The most significant result was that the design center of gravity led to a condition of static instability at small angles of attack at Mach number 17. At this Mach number, the static stability was a highly nonlinear function of the angle of attack. A useful method for analyzing free-flight data having this nonlinear behavior is included in this report. Comparisons were made between the measured aerodynamic coefficients and those estimated by Newtonian impact theory and by a method developed by Seiff and Whiting. The latter method gave good estimates of the normal-force-curve slope at both Mach numbers and of the moment-curve slope at the lower Mach number. It resulted in an overestimation of the static stability at Mach number 17, although it gave results decidedly closer to the experimental value than did Newtonian impact theory.

FREE FLIGHT↗
NASA NTRS1990

The vortex flap F-106B, overcoming safety and data problems in flight testing

A NASA, F-106B airplane equipped with a vortex flap is establishing a data base for calibration of design and analysis tools for this vortical flow control concept. Some of the problems in safely and precisely flying to research data points were caused by the modified airplane's low longitudinal stability and low margin of structural strength. These characteristics are, in part, a consequence of the program's low cost approach which utilizes a bolt-on, ground-adjustable flap attached to a minimally modified airplane. Presently, the 40-degree flap deflection flight research is being concluded with a flow visualization phase. A similar pattern of flight envelope expansion, pressure distribution, performance, and flow visualization phases remain to be accomplished for the 30-degree flap setting. To date, good research flight data has followed resolution of the safety and data problems described.

Brown, Philip W.↗
NASA NTRS1991

GPS interferometric attitude and heading determination - Flight test results

Methods and results for realtime pitch, roll, and heading determination of a dynamic flexible platform are presented. Consideration is given to a GPS attitude and heading determination system installed in Ohio University's Douglas DC-3 research aircraft. Signals from four antennas are processed by an Ashtech 3DF 24-channel GPS receiver. Data from the receiver are sent to a microcomputer for storage and further computations. Attitude and heading data are sent to a second computer for display on a software generated artificial horizon. Demonstration of this technique proves its candidacy for augmentation of aircraft state estimation for flight control and navigation as well as for numerous other applications, such as reliable aircraft state estimation, measurement of structural flexing of large space structures, synthetic aperture radar, aerial photogrammetry, pointing systems, spacecraft attitude control, rendezvous and docking, and road mapping.

Van Graas, Frank↗