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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Experimental measurements and analytical analysis related to gas turbine heat transfer. Part 1: Time-averaged heat-flux and surface-pressure measurements on the vanes and blades of the SSME fuel-side turbine and comparison with prediction. Part 2: Phase-resolved surface-pressure and heat-flux measurements on the first blade of the SSME fuel-side turbine

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

Source record↗
NASA NTRS1998

Rotor Blade Pressure Measurement in a Rotating Machinery Using Pressure and Temperature Sensitive Paints

Pressure and temperature sensitive paints have been utilized for the measurement of blade surface pressure and temperature distributions in a high speed axial compressor and an Allied Signal F109 gas turbine engine. Alternate blades were painted with temperature sensitive paints and then pressure sensitive paint. This combination allows temperature distributions to be accounted for when determining the blade suction surface pressure distribution. Measurements were taken and pressure maps on the suction surface of a blade were obtained over a range of rotational speeds. Pressure maps of the suction surface show-strong shock waves at the higher speeds.

Torgerson, S.↗
NASA NTRSDate not supplied

Adaptive Pressure Profile Method to Locate the Isolator Shock Train Leading Edge Given Limited Pressure Information

To maximize the performance of high-speed air-breathing engines, such as dual-mode scramjets, the streamwise location of the shock train leading edge (STLE) is ideally placed as far upstream in the isolator as possible while avoiding engine unstart. Thus, it is of interest to quantify and control the STLE location as the vehicle travels along its flight trajectory. The STLE location is typically quantified using wall static pressure measurements but there are often restrictions on the number and placement of transducers, thus reducing the accuracy and overall capability of STLE detection methods. In this work, the Adaptive Pressure Profile (APP) method is introduced to address such sparsity concerns. This method is data driven and does not heavily rely on prior information about the flow regime or engine model. Instead, the \mname method uses real-time pressure measurements from a small number of transducers to adaptively learn the isolator pressure profile. This adaptively-learned profile is fit to the pressure data at each time instance to estimate STLE location. The \mname method produces accurate estimates even when (1) the STLE location is not bounded by two or more transducers or (2) when the STLE location is between two transducers that are situated several duct heights apart. Data from two direct-connect isolator models are used to evaluate the accuracy of the \mname method and demonstrate its robustness for different back-pressure scenarios and transducer configurations.

Robin L Hunt↗
NASA NTRSDate not supplied

Adaptive Pressure Profile Method to Locate the Isolator Shock Train Leading Edge Given Limited Pressure Information

To maximize the performance of high-speed air-breathing engines, such as dual-mode scramjets, the streamwise location of the shock train leading edge (STLE) is ideally placed as far upstream in the isolator as possible while avoiding engine unstart. Thus, it is of interest to quantify and control the STLE location as the vehicle travels along its flight trajectory. The STLE location is typically quantified using wall static pressure measurements but there are often restrictions on the number and placement of transducers, thus reducing the accuracy and overall capability of STLE detection methods. In this work, the Adaptive Pressure Profile (APP) method is introduced to address such sparsity concerns. This method is data driven and does not heavily rely on prior information about the flow regime or engine model. Instead, the \mname method uses real-time pressure measurements from a small number of transducers to adaptively learn the isolator pressure profile. This adaptively-learned profile is fit to the pressure data at each time instance to estimate STLE location. The \mname method produces accurate estimates even when (1) the STLE location is not bounded by two or more transducers or (2) when the STLE location is between two transducers that are situated several duct heights apart. Data from two direct-connect isolator models are used to evaluate the accuracy of the \mname method and demonstrate its robustness for different back-pressure scenarios and transducer configurations.

Robin Hunt↗
NASA NTRS2021

Automated Technique for the Detection of Step-Like Solar Wind Dynamic Pressure Changes: Application to the Response of the Transpolar Potential to Solar Wind Dynamic Pressure Fronts

The input from the solar wind (SW) and the accompanying Interplanetary Magnetic Field (IMF) is the defining feature of the solar wind-magnetosphere interactions. One of the most important SW drivers is the solar wind dynamic pressure, and most significantly its sharp changes known as pressure fronts. Sudden enhancements in solar wind dynamic pressure, SWP, have been shown to lead to intensification of the auroral electrojets (and Region 1 field-aligned currents), increase in auroral emissions, and significant enhancement in ionospheric convection and the transpolar potential. Case studies of the interaction between pressure fronts and the terrestrial space environment shed great light on their importance to the structure and dynamics of the system. However, thorough understanding of the interaction requires fully statistical investigations. These statistical studies require a large number of suitable pressure front events. In the past these have been obtained through laborious manual inspection of solar wind data. We present an automated procedure that identifies sudden step enhancements in SWP that satisfy a desired set of limiting step-like characteristics, such as magnitude of the change, sign of the change (positive/negative), and rate of change. The procedure yields a list of suitable events, along with their step features within minutes. The technique is tested for missed events and false positives by direct comparison to manual inspection of the data. Its utility is illustrated by application to the statistical effect of pressure fronts to the transpolar potential as measured by the Assimilative Mapping of Ionospheric Electrodynamics (AMIE) technique.

A. Boudouridis↗
NASA NTRSDate not supplied

Spectral Analysis of Integrated Pressures on Patches with Unsteady Pressure-Sensitive Paint Measurements

This paper describes the spectral analysis of integrated pressures on patches of the scale model of the Space Launch System Block 1 cargo vehicle with the Unsteady Pressure-Sensitive Paint measurements, which were collected in the ascent transient aerodynamics test with the Unitary Plan Wind Tunnel 11-by-11-foot Transonic Wind Tunnel in September 2019 at NASA Ames Research Center. The patches are defined with x station values, indicating the position along length of the vehicle, and azimuth angles of the scale model. For each patch, the polygons are determined from the surface cells of the grid of the model, clipped with the edges of the patch, and each of the polygons is divided into triangles. The integrated pressure of the patch is determined as the ratio of the sum of the forces on the triangles over the sum of the areas of the triangles. For each run of the test, the cross power spectral density and magnitude squared coherence are computed from the time series of the integrated pressures on the patches. The results of pressure integration and spectral analysis are presented, and the data consistency of the test is demonstrated. The work described in this paper is a part of NASA’s development of a new state-of-the-art uPSP capability in production wind tunnels. Funding for this research was provided by the NASA Aeroscience Evaluation and Test Capabilities Project.

Pressure-Sensitive Paint↗
NASA NTRSDate not supplied

Methodology for Validation of Unsteady Pressure-Sensitive Paint Measurements using Pressure Transducers

Unsteady Pressure-Sensitive Paint (uPSP) technology provides simultaneously high-frequency and high-spatial resolution surface pressure fluctuation measurements. It has demonstrated value in wind tunnel test facilities to evaluate aerospace vehicle unsteady aerodynamics. The calibration of measurements from digital pixel counts into corresponding physical units relies on both lab-based and in-situ calibration steps, and its validation is both challenging and critical to adoption of the technology. We present a rigorous methodology for direct comparison of uPSP-derived surface pressure measurements to reference pressure transducer data based on a model of the camera-based measurement error, a model of the flow field pressure signal, and their propagation through the uPSP data reduction algorithm. The methodology is demonstrated on data collected in a small-scale, subsonic wind tunnel test as well as data collected from a Space Launch System unsteady aerodynamics characterization test at the NASA Ames Research Center Unitary Plan Wind Tunnel Complex.

Pressure-Sensitive Paint↗
NASA NTRSDate not supplied

Degradation Characterization of Unsteady Pressure-Sensitive Paint Under Various Pressures

The lifetime and intensity techniques are two distinct methods used to measure the surface pressure of wind tunnel models with pressure-sensitive paint (PSP). When exposed to 405 nm light for extended periods, both the luminescence and the lifetime ratio were observed to decrease, affecting perceived pressure values. To help further the implementation of the uPSP, this paper describes the efforts to characterize the degradation of both the intensity of the paint and its lifetime ratio, as well as the change in intensity calibration. It was found that the paint degrades at different rates depending on the partial pressure of oxygen during excitation. It was also found that the calibration curves exhibit greater changes when under higher oxygen concentrations, both indicating photodegradation due to photodecomposition and photo-oxidation processes. These effects are observed on models in a subsonic 14” × 14” wind tunnel. The work provides insight into the dynamic degradation behavior of PSP and uPSP, paving the way to reduce the error in pressure measurements obtained through these techniques.

Pressure-Sensitive Paint↗
NASA NTRS1971

Miniature pressure valve for on-board calibration of pressure transducers.

This paper describes the development and testing of a miniature one-pole, three-position pressure valve and actuation mechanism. This valve was designed primarily for installation with a pressure transducer directly in a wind-tunnel model, thereby reducing lag time effects resulting from long connecting tubes to the transducer. This valve provides for the in-situ calibration of the transducer during testing by providing one position for transducer zero, a second position for a known calibrate pressure, and a third position for measurement of the unknown test pressure. The valve is driven from the spring loaded test pressure position to the zero and calibrate positions by a pneumatically operated dual-piston yoke assembly. Large numbers of transducers can be simultaneously calibrated from a remote location in a matter of seconds.

Rollins, G. N.↗
NASA NTRS1973

High-pressure combustor exhaust emissions with improved air-atomizing and conventional pressure-atomizing fuel nozzles

A high-pressure combustor segment 0.456 meter (18 in.) long with a maximum cross section of 0.153 by 0.305 meter (6 by 12 in.) was tested with specially designed air-atomizing and conventional pressure-atomizing fuel nozzles at inlet-air temperatures of 340 to 755 k (610 deg to 1360 R), reference velocities of 12.4 to 26.1 meters per second (41 to 86 ft/sec), and fuel-air ratios of 0.008 to 0.020. Increasing inlet-air pressure from 4 to 20 atmospheres generally increased smoke number and nitric oxide, but decreased carbon monoxide and unburned hydrocarbon concentrations with air-atomizing and pressure-atomizing nozzles. Emission indexes for carbon monoxide and unburned hydrocarbons were lower at 4, 10, and 20 atmospheres, and nitric oxide emission indexes were lower at 10 and 20 atmospheres with air-atomizing than with pressure-atomizing nozzles.

Ingebo, R. D.↗
NASA NTRS1976

A dynamic pressure source for the calibration of pressure transducers

A dynamic pressure source is described for producing sinusoidally varying pressures of up to 34 kPa zero to peak, over the frequency range of approximately 50 Hz to 2 kHz. The source is intended for the dynamic calibration of pressure transducers. The transducer to be calibrated is mounted near the base of the thick walled aluminum tube forming the vessel so that the pressure sensitive element is in contact with the liquid in the tube. A section of the tube is filled with small steel balls to damp the motion of the 10-St dimethyl siloxane working fluid in order to extend the useful frquency range to higher frequencies than would be provided by an undamped system. The dynamic response of six transducers provided by the sponsor was evaluated using the pressure sources; the results of these calibrations are given.

Vezzetti, C. F.↗
NASA NTRS1987

Some features of surface pressure fluctuations in turbulent boundary layers with zero and favorable pressure gradients

Measurements of surface pressure fluctuation spectra, coherence and convective wave speeds from zero and favorable pressure gradient turbulent boundary layers are reported for momentum Reynolds numbers from 3000 to 18,800. The acceleration parameter K is near 2 x 10 to the -7 power for the favorable pressure gradient flow. The outer variables, U sub e, tau sub w and delta sub 1 non-dimensionalize and collapse the spectra for the low to middle range of frequencies for most test cases. The grouping using the inner variable, U sub tau and gamma, collapse the spectra for the middle to high range of frequencies for all test cases. The value of p'/tau sub w was near 3.8 and 2.8 for the smallest values of d+ in the zero and favorable pressure gradient flows, respectively. The coherence exhibits a decay that is not exponential in some cases, but the Corcos similarity parameters omega Delta x/U sub c and omega Delta z/U sub c collapse the data for all test cases. The ratio of U sub c/U sub e increases with omega delta sub 1/U sub e up to omega delta sub 1/U sub e on the order of unity, where U sub c/U sub e becomes nearly constant. This was observed in the present results for both streamwise pressure gradient flows. The experimental results presented show good agreement with previous research.

Mcgrath, B. E.↗
NASA NTRS1987

Mechanical characteristics of filament-wound pressure vessel (burst pressure)

The finite element method is used to analyze the mechanical characteristics of a pressurized filament-wound (FW) pressure vessel, and to predict its burst pressure. The analysis takes into account the bending moment, the stretch-bend coupling effect, nonlinear stress-strain relations, and finite deflection. The analysis is based on two initial failure criteria for laminae, and two ultimate fracture criteria for laminated structures. The numerical results, obtained by applying the load incremental method to the isotensoid CFRP pressure vessel used in the launching of the Zikiken satellite, are in good agreement with the experimental burst pressure and fracture behaviors.

Iida, H.↗
NASA NTRS1990

Aerodynamic pressure and heating-rate distributions in tile gaps around chine regions with pressure gradients at a Mach number of 6.6

Surface and gap pressures and heating-rate distributions were obtained for simulated Thermal Protection System (TPS) tile arrays on the curved surface test apparatus of the Langley 8-Foot High Temperature Tunnel at Mach 6.6. The results indicated that the chine gap pressures varied inversely with gap width because larger gap widths allowed greater venting from the gap to the lower model side pressures. Lower gap pressures caused greater flow ingress from the surface and increased gap heating. Generally, gap heating was greater in the longitudinal gaps than in the circumferential gaps. Gap heating decreased with increasing gap depth. Circumferential gap heating at the mid-depth was generally less than about 10 percent of the external surface value. Gap heating was most severe at local T-gap junctions and tile-to-tile forward-facing steps that caused the greatest heating from flow impingement. The use of flow stoppers at discrete locations reduced heating from flow impingement. The use of flow stoppers at discrete locations reduced heating in most gaps but increased heating in others. Limited use of flow stoppers or gap filler in longitudinal gaps could reduce gap heating in open circumferential gaps in regions of high surface pressure gradients.

Hunt, L. Roane↗
NASA NTRS1990

Determination of the Pressure Drag of Airfoils by Integration of Surface Pressures

A study was conducted of the causes of pressure drag of subsonic airfoils. In a previous paper by the author, the pressure drag is obtained by calculating the total drag from the momentum defect in the boundary layer at the trailing edge and subtracting the friction drag obtained from integration of surface friction along the chord. Herein, the pressure drag is obtained by integrating the streamwise components of surface pressure around the airfoil. Studies were made to verify the accuracy of the integration procedure. The values of pressure drag were much smaller than those obtained by the previous method. This lack of agreement is attributed to the difficulty of calculating boundary layer conditions in the vicinity of the trailing edge and to the extreme sensitivity of the circulation and lift to the trailing edge conditions. The results of these studies are compared with those of previous investigations.

Phillips, William H.↗
NASA NTRS1989

A temperature and pressure controlled calibration system for pressure sensors

A data acquisition and experiment control system capable of simulating temperatures from -184 to +220 C and pressures either absolute or differential from 0 to 344.74 kPa is developed to characterize silicon pressure sensor response to temperature and pressure. System software is described that includes sensor data acquisition, algorithms for numerically derived thermal offset and sensitivity correction, and operation of the environmental chamber and pressure standard. This system is shown to be capable of computer interfaced cryogenic testing to within 1 C and 34.47 Pa of single channel or multiplexed arrays of silicon pressure sensors.

Chapman, John J.↗
NASA NTRS1992

Pressure oscillation in the leakage annulus between a shrouded impeller and its housing due to impeller-discharge-pressure disturbances

The perturbed flow in the leakage path between a shrouded-pump impeller and its housing is analyzed using experiences with the Space Shuttle Main Engine (SSME), high pressure fuel turbopump (HPFTP) wearing-ring seals. Analysis is based on a bulk-flow model which consists of the path-momentum, circumferential momentum, and continuity equations. The pressure oscillations in the leakage annulus are driven by a circumferential variation of the impeller discharge pressure. It is shown that the occurrence and nature of the pressure oscillations depend on the tangential-velocity ratio of the fluid entering the seal, the order of the Fourier coefficient, the closeness of the precessional frequency of the rotating pressure field to the first natural frequency of the fluid annulus, and the clearance of the wearing-ring seal. The results obtained may explain the internal melting observed on SSME HPFTP seal parts.

Childs, D. W.↗
NASA NTRS1994

Non-invasive method and apparatus for monitoring intracranial pressure and pressure volume index in humans

Non-invasive measuring devices responsive to changes in a patient's intracranial pressure (ICP) can be accurately calibrated for monitoring purposes by providing known changes in ICP by non-invasive methods, such as placing the patient on a tilting bed and calculating a change in ICP from the tilt angle and the length of the patient's cerebrospinal column, or by placing a pressurized skull cap on the patient and measuring the inflation pressure. Absolute values for the patient's pressure-volume index (PVI) and the steady state ICP can then be determined by inducing two known changes in the volume of cerebrospinal fluid while recording the corresponding changes in ICP by means of the calibrated measuring device. The two pairs of data for pressure change and volume change are entered into an equation developed from an equation describing the relationship between ICP and cerebrospinal fluid volume. PVI and steady state ICP are then determined by solving the equation. Methods for inducing known changes in cerebrospinal fluid volume are described.

Cantrell, John H.↗