Results for “Pressure transducers” — Page 17

NASA NTRS2021

Hypersonic Second-Mode Instability Response to Shaped Roughness

An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.

Hypersonics↗

NASA NTRS2021

Hypersonic Second-Mode Instability Response to Shaped Roughness

An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.

Hypersonics↗

NASA NTRS2021

Hypersonic Second Mode Instability Response to Shaped Roughness

An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.

Hypersonics↗

NASA NTRS1971

System accurately controls pressure in cryogenic tanks

High-resolution differential pressure transducer senses very small positive or negative pressure variations in the cryogenic tank relative to absolute reference pressure. System is useful in calibrating instruments where working fluid must be maintained at closely controlled temperature, or in processes requiring very fine pressure control.

Kirchmeier, W. E., Jr.↗

NASA NTRS1974

Environmental control system transducer development study

A development test program of transducers for aerospace projects is described. Stability and performance of existing transducers, and improvements compatible with shuttle ECS requirements are investigated. These requirements incorporate design and development features into the transducers, and include the following: (1) improvement of overall transducer ruggedness and reliability; (2) common transducers for all ECS fluids that will be unaffected by long quiescent periods in the space environment, that will require no maintenance or refurbishing for at least 100 launches; and (3) appropriate self-check features that simplify checkout and maintenance. Models of three different transducers, a three-way valve for pressure transducers from closed liquid loops, surface-type platinum-wire resistance temperature sensors, and a nuclenics gaging system are evaluated. Tests and development improvements are described.

Brudnicki, M. J.↗

NASA NTRS1982

Fluctuating pressures on fan blades of a turbofan engine: Static and wind-tunnel investigations

To investigate the fan noise generated from turbofan engines, miniature pressure transducers were used to measure the fluctuating pressure on the fan blades of a JT15D engine. Tests were conducted with the engine operating on an outdoor test stand and in a wind tunnel. It was found that a potential flow interaction between the fan blades and six, large support struts in the bypass duct is a dominant noise source in the JT15D engine. Effects of varying fan speed and the forward speed on the blade pressure are also presented.

Schoenster, J. A.↗

NASA NTRS1985

Detailed pressure distribution measurements obtained on several configurations of an aspect-ratio-7 variable twist wing

Detailed pressure distribution measurements were made for 11 twist configurations of a unique, multisegmented wing model having an aspect ratio of 7 and a taper ratio of 1. These configurations encompassed span loads ranging from that of an untwisted wing to simple flapped wings both with and without upper-surface spoilers attached. For each of the wing twist configurations, electronic scanning pressure transducers were used to obtain 580 surface pressure measurements over the wing in about 0.1 sec. Integrated pressure distribution measurements compared favorably with force-balance measurements of lift on the model when the model centerbody lift was included. Complete plots and tabulations of the pressure distribution data for each wing twist configuration are provided.

Holbrook, G. T.↗

NASA NTRS1985

Fluctuating pressures on fan blades of a turbofan engine, flight test investigation

Miniature pressure transducers were used to measure the fluctuating pressure on the fan blades of a JT15D engine in flight to aid in understanding fan-noise generation. Although the blade pressures can be considered very useful in determining far-field noise sources in the subsonic fan-blade tip speeds, other mechanisms, and as shock waves in the transonic and supersonic fan-blade tip speeds, limit their usefulness in these higher speed ranges.

Schoenster, J. A.↗

NASA NTRS1986

Dynamic Pressure Calibration Standard

Vibrating columns of fluid used to calibrate transducers. Dynamic pressure calibration standard developed for calibrating flush diaphragm-mounted pressure transducers. Pressures up to 20 kPa (3 psi) accurately generated over frequency range of 50 to 1,800 Hz. System includes two conically shaped aluminum columns one 5 cm (2 in.) high for low pressures and another 11 cm (4.3 in.) high for higher pressures, each filled with viscous fluid. Each column mounted on armature of vibration exciter, which imparts sinusoidally varying acceleration to fluid column. Signal noise low, and waveform highly dependent on quality of drive signal in vibration exciter.

Schutte, P. C.↗

NASA NTRS1987

University of Missouri-Rolla cloud simulation facility - Proto II chamber

The design and supporting systems for the cooled-wall expansion cloud chamber, designated Proto II, are described. The chamber is a 10-sided vertical cylinder designed to be operated with interior wall temperatures between +40 and -40 C, and is to be utilized to study microphysical processes active in atmospheric clouds and fogs. Temperatures are measured using transistor thermometers which have a range of + or - 50 C and a resolution of about + or - 0.001 C; and pressures are measured in the chamber by a differential strain gauge pressure transducer. The methods used for temperature and pressure control are discussed. Consideration is given to the chamber windows, optical table, photographic/video, optical attenuation, Mie scattering, and the scanning system for the chamber. The system's minicomputer and humidifier, sample preparation, and chamber flushing are examined.

White, Daniel R.↗

NASA NTRS1990

Air-Velocity Sensor For Helicopter

New airspeed sensor conceived for accurate measurement of both airspeed and direction of flight of helicopter. Direction of motion of helicopter displayed by lighting of one of series of lamps encircling digital display of airspeed. Pressure transducer measures difference between impact and static pressures at tip of rotor blade by use of conventional pitot-static-tube assembly.

Garner, H. Douglas↗

NASA NTRS1989

Non-contact temperature measurements in support of microgravity combustion experiments

Recent conceptual advances in the understanding of combustion science fundamentals in the context of microgravity processes and phenomenology have resulted in an increased demand for diagnostic systems of greater sophistication. Owing primarily to the severe operational constraints that accompany the space flight environment, measurement systems to date remain fairly primative in nature. Qualitative pictures provided by photographic recording media comprise the majority of the existing data, the remainder consisting of the output of conventional transducers, such as thermocouples, hot wires, and pressure transducers. The absence of the rather strong influence of buoyant convection renders microgravity combustion phenomena more fragile than their 1-G counterparts. The emphasis was placed on nonperturbing optical diagnostics. Other factors such as limited supplies of expendable reactants, and periods of microgravity time of sufficient duration, coupled with more fundamental questions regarding inherent length and time scales and reproducibility have favored multipoint or multidimensional techniques. While the development of optical diagnostics for application to combustion science is an extremely active area at present, the peculiarities of space flight hardware severely restrict the feasibility of implementing the majority of techniques which are being utilized in terrestrial applications. The additional requirements for system reliability and operational simplicity have tended to promote somewhat less commonly emphasized techniques such as refractive index mapping and molecular Rayleigh scattering, which are briefly discussed.

Greenberg, Paul S.↗

NASA NTRS1989

Pressure Distribution and Air Data System for the Aeroassist Flight Experiment

The Aeroassist Flight Experiment (AFE) is designed to provide critical flight data necessary for the design of future Aeroassist Space Transfer Vehicles (ASTV). This flight experiment will provide aerodynamic, aerothermodynamic, and environmental data for verification of experimental and computational flow field techniques. The Pressure Distribution and Air Data System (PD/ADS), one of the measurement systems incorporated into the AFE spacecraft, is designed to provide accurate pressure measurements on the windward surface of the vehicle. These measurements will be used to determine the pressure distribution and air data parameters (angle of attack, angle of sideslip, and free-stream dynamic pressure) encountered by the blunt-bodied vehicle over an altitude range of 76.2 km to 94.5 km. Design and development data are presented and include: measurement requirements, measurement heritage, theoretical studies to define the vehicle environment, flush-mounted orifice configuration, pressure transducer selection and performance evaluation data, and pressure tubing response analysis.

Gibson, Lorelei S.↗

NASA NTRS1992

Time-averaged heat transfer and pressure measurements and comparison with prediction for a two-stage turbine

Time-averaged Stanton number and surface-pressure distributions are reported for the first-stage vane row and the first-stage blade row of the Rocketdyne Space Shuttle Main Engine two-stage fuel-side turbine. These measurements were made at 10 percent, 50 percent, and 90 percent span on both the pressure and suction surfaces of the component. Stanton-number distributions are also reported for the second-stage vane at 50 percent span. A shock tube is used as a short-duration source of heated and pressurized air to which the turbine is subjected. Platinum thin-film gages are used to obtain the heat-flux measurements and miniature silicone-diaphragm pressure transducers are used to obtain the surface pressure measurements. The first-stage vane Stanton number distributions are compared with predictions obtained using a quasi-3D Navier-Stokes solution and a version of STAN5. This same N-S technique was also used to obtain predictions for the first blade and the second vane.

Dunn, M. G.↗

NASA NTRS1995

Measurement of Gust Response on a Turbine Cascade

The paper presents benchmark experimental data on a gust response of an annular turbine cascade. The experiment was particularly designed to provide data for comparison with the results of a typical linearized gust-response analysis. Reduced frequency, Mach number, and incidence were varied independently. Except for the lowest reduced frequency, the gust velocity distribution was nearly sinusoidal. For the high inlet-velocity series of tests, the cascade was near choking. The mean flow was documented by measuring blade surface pressures and the cascade exit flow. High-response pressure transducers were used to measure the unsteady pressure distribution. Inlet-velocity components and turbulence parameters were measured using hot wire. In addition to the synchronous time-average pressure spectra, typical power spectra are included for several representative conditions.

Kurkov, A. P.↗

NASA NTRS1995

Fundamental phenomena on fuel decomposition and boundary-layer combustion processes with applications to hybrid rocket motors

The experimental study on the fundamental processes involved in fuel decomposition and boundary-layer combustion in hybrid rocket motors is continuously being conducted at the High Pressure Combustion Laboratory of The Pennsylvania State University. This research will provide a useful engineering technology base in the development of hybrid rocket motors as well as a fundamental understanding of the complex processes involved in hybrid propulsion. A high-pressure, 2-D slab motor has been designed, manufactured, and utilized for conducting seven test firings using HTPB fuel processed at PSU. A total of 20 fuel slabs have been received from the Mcdonnell Douglas Aerospace Corporation. Ten of these fuel slabs contain an array of fine-wire thermocouples for measuring solid fuel surface and subsurface temperatures. Diagnostic instrumentation used in the test include high-frequency pressure transducers for measuring static and dynamic motor pressures and fine-wire thermocouples for measuring solid fuel surface and subsurface temperatures. The ultrasonic pulse-echo technique as well as a real-time x-ray radiography system have been used to obtain independent measurements of instantaneous solid fuel regression rates.

Kuo, Kenneth K.↗

NASA NTRS1996

Apparatus Tests Peeling Of Bonded Rubbery Material

Instrumented hydraulic constrained blister-peel apparatus obtains data on degree of bonding between specimen of rubbery material and rigid plate. Growth of blister tracked by video camera, digital clock, pressure transducer, and piston-displacement sensor. Cylinder pressure controlled by hydraulic actuator system. Linear variable-differential transformer (LVDT) and float provide second, independent measure of change in blister volume used as more precise volume feedback in low-growth-rate test.

Crook, Russell A.↗

NASA NTRS2001

Investigation of Capillary Limit in a Loop Heat Pipe

This paper presets an experimental study on the capillary limit of a loop heat pipe (LHP) at low powers. The slow thermal response of the loop at low powers made it possible to observe interactions among various components after the capillary limit was exceeded. The capillary limit at low powers was achieved by imposing additional pressure drops on the vapor line through the use of a metering valve. A differential pressure transducer was also used to measure the pressure drop across the evaporator and the compensation chamber (CC). Test results show that when the capillary limit is exceeded, vapor will penetrate the primary wick, resulting in a partial dry-out of the evaporator and a rapid increase of the CC temperature. Because the evaporator can tolerate vapor bubbles, the LHP will continue to function and may reach a new steady state at the higher temperature. Thus, the LHP will exhibit a graceful degradation in performance rather than a complete failure. Moreover, the loop can recover from a partial dry-out by reducing the heat load without a re-start.

Ku, Jentung↗