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  • 1
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    Design and Laboratory Validation of a Capacitive Sensor for Measuring the Recession of Thin-Layered Ablator (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Flight vehicles are typically instrumented with subsurface thermocouples to estimate heat transfer at the surface using inverse analysis procedures. If the vehicle has an ablating heat shield, however, temperature time histories from subsurface thermocouples no longer provide enough information to estimate heat flux at the surface. In this situation, the geometry changes and thermal energy leaves the surface in the form of ablation products. The ablation rate is required to estimate heat transfer to the surface. A new concept for a capacitive sensor has been developed to measure ablator depth using the ablator's dielectric effect on a capacitor's fringe region. Relying on the capacitor's fringe region enables the gage to be flush mounted in the vehicle's permanent structure and not intrude into the ablative heat shield applied over the gage. This sensor's design allows nonintrusive measurement of the thickness of dielectric materials, in particular, the recession rates of low-temperature ablators applied in thin (0.020 to 0.060 in. (0.05 to 0.15 mm)) layers. Twenty capacitive gages with 13 different sensing element geometries were designed, fabricated, and tested. A two-dimensional finite-element analysis was performed on several candidate geometries. Calibration procedures using ablator-simulating shims are described. A one-to-one correspondence between system output and dielectric material thickness was observed out to a thickness of 0.055 in. (1.4 mm) for a material with a permittivity about three times that of air or vacuum. A novel method of monitoring the change in sensor capacitance was developed. This technical memorandum suggests further improvements in gage design and fabrication techniques.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TM-4777 , H-2111 , NAS 1.15:4777
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Aerothermal test results from the second flight of the Pegasus booster (1992)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A survey of temperature, heat-flux, and pressure measurements was obtained at speeds through Mach 8.0 on the second flight of the Pegasus air-launched space booster system. All sensors were distributed on the wing-body fairing or fillet. Sensors included thin foil-gauge thermocouples installed near the surface within the thermal protection system. Thermocouples were also installed on the surface of nonablating plugs. The resulting temperature time history allowed derivation of convective heat flux. In addition, commercially available calorimeters were installed on the fillet at selected locations. Calorimeters exhibited a larger change in measured heat flux than collocated nonablating plugs in response to particular events. Similar proportional variations in heat flux across different regions of the fillet were detected by both the calorimeters and nonablating plugs. Pressure ports were installed on some nonablating plugs to explore the effects of port protrusion and high-frequency noise on pressure requirements. The effect of port protrusion on static-pressure measurements was found to decrease with increasing Mach number. High-frequency noise suppression was found to be desirable but not required on any future flight.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-4391 , H-1827 , NAS 1.15:4391
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    NASA TECHNICAL REPORTS
  • 3
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    Summary of aerothermal test results from the first flight of the Pegasus air-launched space booster (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Temperature measurements were obtained on the Pegasus booster from launch through Mach 8.0. The majority of sensors were thin-foil temperature gages installed near the surface within the vehicle's ablating thermal protection system. These gages were distributed on the wing surfaces and on the wing-body fairing or fillet. Temperature time histories from these installations are presented. In addition, thermocouples were installed on the surface of nonablating plugs located on the fairing. These sensors were more responsive to changes in flight conditions than the foil gages and allowed a derivation of convective heat flux. A heating rate magnification of 2 was found in the vicinity of the wing shock interaction.
    Keywords: LAUNCH VEHICLES AND SPACE VEHICLES
    Type: AIAA PAPER 91-5046
    Format: text
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  • 4
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    Aerothermal test results from the first flight of the Pegasus air-launched space booster (1991)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A survey of temperature measurements at speeds through Mach 8.0 on the first flight of the Pegasus air-launched booster system is discussed. In addition, heating rates were derived from the temperature data obtained on the fuselage in the vicinity of the wing shock interaction. Sensors were distributed on the wing surfaces, leading edge, and on the wing-body fairing or fillet. Side-by-side evaluations were obtained for a variety of sensor installations. Details of the trajectory reconstruction through first-stage separation are provided. Given here are indepth descriptions of the sensor installations, temperature measurements, and derived heating rates along with interpretations of the results.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-4330 , H-1672 , NAS 1.15:4330
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  • 5
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    Experimental Evaluation of Hot Films on Ceramic Substrates for Skin-Friction Measurement (2003)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An investigation has been performed on the use of low-thermal conductivity, ceramic substrates for hot films intended to measure skin friction. Hot films were deposited on two types of ceramic substrates. Four hot films used composite-ceramic substrates with subsurface thermocouples (TCs), and two hot films were deposited on thin Macor(R) substrates. All six sensors were tested side by side in the wall of the NASA Glenn Research Center 8-ft by 6-ft Supersonic Wind Tunnel (SWT). Data were obtained from zero flow to Mach 1.98 in air. Control measurements were made with three Preston tubes and two boundary-layer rakes. The tests were repeated at two different hot film power levels. All hot films and subsurface TCs functioned throughout the three days of testing. At zero flow, the films on the high-thermal conductivity Macor(R) substrates required approximately twice the power as those on the composite-ceramic substrates. Skin-friction results were consistent with the control measurements. Estimates of the conduction heat losses were made using the embedded TCs but were hampered by variability in coating thicknesses and TC locations.
    Keywords: Nonmetallic Materials
    Type: NASA/TM-2003-210742 , NAS 1.15:210742 , H-2517 , 6th ASME-JSME Thermal Engineering Joint Conference; Mar 16, 2003 - Mar 20, 2003; Hawaii Island, HI; United States
    Format: text
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    NASA TECHNICAL REPORTS
  • 6
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    Exploratory Calibration of Adjustable-Protrusion Surface-Obstacle (APSO) Skin Friction Vector Gage (2003)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The design of an adjustable-protrusion surface-obstacle (APSO) skin friction vector gage is presented. Results from exploratory calibrations conducted in laminar and turbulent boundary layers at the Washington University Low-Speed Wind Tunnel and for turbulent boundary layers at speeds up to Mach 2 on the ceiling of the NASA Glenn Research Center 8- X 6-ft Supersonic Wind Tunnel are also discussed. The adjustable-height gage was designed to yield both the magnitude and direction of the surface shear stress vector and to measure the local static pressure distribution. Results from the NASA test show good correlation for subsonic and low supersonic conditions covering several orders of magnitude in terms of the adopted similarity variables. Recommendations for future work in this area consist of identifying the physical parameters responsible for the disagreement between the university and NASA data sets, developing a compressibility correction specific to the APSO geometry, and examining the effect that static pressure distribution and skewed boundary layers have on the results from the APSO.
    Keywords: Instrumentation and Photography
    Type: NASA/TM-2003-210739 , H-2519 , NAS 1.15:210739 , AIAA Paper 2003-0740 , 41st AIAA Aerospace Sciences Meeting and Exhibit; Jan 06, 2003 - Jan 09, 2003; Reno, NV; United States
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    NASA TECHNICAL REPORTS
  • 7
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    Pegasus(Registered trademark) Wing-Glove Experiment to Document Hypersonic Crossflow Transition: Measurement System and Selected Flight Results (2000)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In a recent flight experiment to study hypersonic crossflow transition, boundary layer characteristics were documented. A smooth steel glove was mounted on the first stage delta wing of Orbital Sciences Corporation's Pegasus (R) launch vehicle and was flown at speeds of up to Mach 8 and altitudes of up to 250,000 ft. The wing-glove experiment was flown as a secondary payload off the coast of Florida in October 1998. This paper describes the measurement system developed. Samples of the results obtained for different parts of the trajectory are included to show the characteristics and quality of the data. Thermocouples and pressure sensors (including Preston tubes, Stanton tubes, and a "probeless" pressure rake showing boundary layer profiles) measured the time-averaged flow. Surface hot-films and high-frequency pressure transducers measured flow dynamics. Because the vehicle was not recoverable, it was necessary to design a system for real-time onboard processing and transmission. Onboard processing included spectral averaging. The quality and consistency of data obtained was good and met the experiment requirements.
    Keywords: Avionics and Aircraft Instrumentation
    Type: NASA/TM-2000-209016 , NAS 1.15:209016 , H-2395 , AIAA Paper 2000-0505 , 38th Aerospace Sciences; Jan 10, 2000 - Jan 13, 2000; Reno, NV; United States
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  • 8
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    Determination of Sun Angles for Observations of Shock Waves on a Transport Aircraft (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Wing compression shock shadowgraphs were observed on two flights during banked turns of an L-1011 aircraft at a Mach number of 0.85 and an altitude of 35,000 ft (10,700 m). Photos and video recording of the shadowgraphs were taken during the flights to document the shadowgraphs. Bright sunlight on the aircraft was required. The time of day, aircraft position, speed and attitudes were recorded to determine the sun azimuth and elevation relative to the wing quarter chord-line when the shadowgraphs were visible. Sun elevation and azimuth angles were documented for which the wing compression shock shadowgraphs were visible. The shadowgraph was observed for high to low elevation angles relative to the wing, but for best results high sun angles relative to the wing are desired. The procedures and equations to determine the sun azimuth and elevation angle with respect to the quarter chord-line is included in the Appendix.
    Keywords: Aerodynamics
    Type: NASA/TM-1998-206551 , H-2251 , NAS 1.15:206551 , Flow Visualization; Sep 01, 1998 - Sep 04, 1998; Sorrento; Italy
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  • 9
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    Results From F-18B Stability and Control Parameter Estimation Flight Tests at High Dynamic Pressures (2000)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A maximum-likelihood output-error parameter estimation technique has been used to obtain stability and control derivatives for the NASA F-18B Systems Research Aircraft. This work has been performed to support flight testing of the active aeroelastic wing (AAW) F-18A project. The goal of this research is to obtain baseline F-18 stability and control derivatives that will form the foundation of the aerodynamic model for the AAW aircraft configuration. Flight data have been obtained at Mach numbers between 0.85 and 1.30 and at dynamic pressures ranging between 600 and 1500 lbf/sq ft. At each test condition, longitudinal and lateral-directional doublets have been performed using an automated onboard excitation system. The doublet maneuver consists of a series of single-surface inputs so that individual control-surface motions cannot be correlated with other control-surface motions. Flight test results have shown that several stability and control derivatives are significantly different than prescribed by the F-18B aerodynamic model. This report defines the parameter estimation technique used, presents stability and control derivative results, compares the results with predictions based on the current F-18B aerodynamic model, and shows improvements to the nonlinear simulation using updated derivatives from this research.
    Keywords: Aircraft Stability and Control
    Type: NASA/TP-2000-209033 , H-2424 , NAS 1.60:209033
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  • 10
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    Multiaxis thrust vectoring using axisymmetric nozzles and postexit vanes on an F/A-18 configuration vehicle (1991)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A ground-based investigation was conducted on an operational system of multiaxis thrust vectoring using postexit vanes around an axisymmetric nozzle. This thrust vectoring system will be tested on the NASA F/A-18 High Alpha Research Vehicle (HARV) aircraft. The system provides thrust vectoring capability in both pitch and yaw. Ground based data were gathered from two separate tests at NASA Langley Research Center. The first was a static test in the 16-foot Transonic Tunnel Cold-Jet Facility with a 14.25 percent scale model of the axisymmetric nozzle and the postexit vanes. The second test was conducted in the 30 by 60 foot wind tunnel with a 16 percent F/A-18 complete configuration model. Data from the two sets are being used to develop models of jet plume deflection and thrust loss as a function of vane deflection. In addition, an aerodynamic interaction model based on plume deflection angles will be developed. Results from the scale model nozzle test showed that increased vane deflection caused exhaust plume turning. Aerodynamic interaction effects consisted primarily of favorable interaction of moments and unfavorable interaction of forces caused by the vectored jet plume.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101741 , H-1705 , NAS 1.15:101741 , High-Angle-of Attack Technology Conference; Oct 30, 1990 - Nov 01, 1990; Hampton, VA; United States
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