Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Dynamic Pressure”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 469 records · Page 26
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

Empirical model of the interaction of the solar wind with the daytime ionosphere of Venus

Interplanetary probe data on the interaction of the solar wind with the ionosphere of Venus are examined. It is shown that the ionospheric pressure above the ionopause is approximately equal to the dynamic pressure of the solar wind, while the value of magnetic pressure at the ionopause amounts to about 2/3 of the value of external and ionospheric pressures. A first-approximation model is proposed for changes of pressure at the ionopause as a function of height and solar zenith angle. This model makes it possible to determine the 'instantaneous' profile of the ionopause for a given solar-wind pressure.

Vaisberg, O. L.↗
NASA NTRS1985

Computer program for calculating flow parameters and power requirements for cryogenic wind tunnels

A computer program has been written that performs the flow parameter calculations for cryogenic wind tunnels which use nitrogen as a test gas. The flow parameters calculated include static pressure, static temperature, compressibility factor, ratio of specific heats, dynamic viscosity, total and static density, velocity, dynamic pressure, mass-flow rate, and Reynolds number. Simplifying assumptions have been made so that the calculations of Reynolds number, as well as the other flow parameters can be made on relatively small desktop digital computers. The program, which also includes various power calculations, has been developed to the point where it has become a very useful tool for the users and possible future designers of fan-driven continuous-flow cryogenic wind tunnels.

Dress, D. A.↗
NASA NTRS1989

Fluctuating pressures measured beneath a high-temperature, turbulent boundary layer on a flat plate at Mach number of 5

Fluctuating pressures were measured beneath a Mach 5, turbulent boundary layer on a flat plate with an array of piezoresistive sensors. The data were obtained with a digital signal acquisition system during a test run of 4 seconds. Data sampling rate was such that frequency analysis up to 62.5 kHz could be performed. To assess in situ frequency response of the sensors, a specially designed waveguide calibration system was employed to measure transfer functions of all sensors and related instrumentation. Pressure time histories were approximated well by a Gaussian prohibiting distribution. Pressure spectra were very repeatable over the array span of 76 mm. Total rms pressures ranged from 0.0017 to 0.0046 of the freestream dynamic pressure. Streamwise, space-time correlations exhibited expected decaying behavior of a turbulence generated pressure field. Average convection speed was 0.87 of freestream velocity. The trendless behavior with sensor separation indicated possible systematic errors.

Parrott, Tony L.↗
NASA NTRS1959

Flight Behavior of the X-2 Research Airplane to a Mach Number of 3.20 and a Geometric Altitude of 126,200 Feet

The maximum Mach number and altitude capabilities of the Bell X-2 research airplane were achieved during a program conducted by the U.S. Air Force with Bell Aircraft Corp. providing operational support and the National Aeronautics and Space Administration providing instrumentation and advisory engineering assistance. A maximum geometric altitude of 126,200 feet was attained at a static pressure of 9.4 pounds per square foot and a dynamic pressure of 19.1 pounds per square foot. During the last flight of the airplane, a maximum Mach number of 3.20 was reached. The directionally divergent maneuver which terminated the final high Mach number flight was precipitated by the loss in directional stability that resulted from increasing the angle of attack. The yawing moment from the lateral control was sufficient to initiate the divergence and also to cause,, indirectly, rolling moments that were greater than the aileron capabilities of the airplane. The ensuing violent motions-resulting from inertial roll coupling caused the loss of the aircraft.

Day, Richard E.↗
NASA NTRS1950

High-Subsonic Performance Characteristics and Boundary-Layer Investigations of a 12 10-Inch-Inlet-Diameter Conical Diffuser

Performance and boundary-layer data were taken in a 12 degree 10-inch inlet-diameter conical diffuser of 2:1 exit- to inlet-area ratio. These data were taken for two inlet-boundary-layer conditions. The first condition was that of a thinner inlet boundary later (boundary-layer displacement thickness, delta* approximately equal to 0.034) produced by an inlet section approximately 1 inlet diameter in length between the entrance bell and the diffuser. The second condition was a thicker inlet boundary layer (delta* approximately equal to 0.120) produced by an additional inlet section length of approximately 6 diameters. Longitudinal static-pressure distributions were measured fro wall static orifices. Transverse total- and static-pressure surveys were made at the inlet and exit stations. Boundary-layer velocity distributions were measured at seven stations between the inlet and exit. These data were obtained for a Reynolds number (based on inlet diameter) range of 1 x 10(exp 6) to 3.9 x 10(exp 6). The corresponding Mach number range was from M = 0.2 to choking. At the maximum-power-available condition supersonic flow was obtained as far as 4.5 inches downstream from the diffuser inlet with a maximum Mach number of M approximately equal to 1.5. The total-pressure loss through the diffuser in percentage of inlet dynamic pressure was approximately 2.5 percent for the thinner inlet boundary later and 5.5 percent for the thicker inlet boundary later over the lower subsonic range. These valued increased with increasing flow rate- the values for the thicker inlet boundary later more than those for the thinner inlet boundary layer. The diffuser effectiveness, expressed as the ratio of the actual static-pressure rise to the ideal static-pressure rise, was about 85 percent for the thinner inlet boundary layer and about 67 percent for the thicker inlet boundary later in the lower subsonic range. These values decrease with increasing flow rate. Separated flow was observed for both inlet-boundary-layer conditions in the region of adverse pressure gradient just downstream of the transition curvature from inlet section to diffuser. The flow for the thinner-inlet-boundary-layer condition did not fully re-establish itself along the diffuser walls. The thicker inlet-boundary-layer flow, while not completely re-establishing the normal flow pattern downstream of the separated region, did re-establish more successfully than the thinner inlet boundary layer.

Little, B. H., Jr.↗
NASA NTRS2013

Mars Entry Atmospheric Data System Trajectory Reconstruction Algorithms and Flight Results

The Mars Entry Atmospheric Data System is a part of the Mars Science Laboratory, Entry, Descent, and Landing Instrumentation project. These sensors are a system of seven pressure transducers linked to ports on the entry vehicle forebody to record the pressure distribution during atmospheric entry. These measured surface pressures are used to generate estimates of atmospheric quantities based on modeled surface pressure distributions. Specifically, angle of attack, angle of sideslip, dynamic pressure, Mach number, and freestream atmospheric properties are reconstructed from the measured pressures. Such data allows for the aerodynamics to become decoupled from the assumed atmospheric properties, allowing for enhanced trajectory reconstruction and performance analysis as well as an aerodynamic reconstruction, which has not been possible in past Mars entry reconstructions. This paper provides details of the data processing algorithms that are utilized for this purpose. The data processing algorithms include two approaches that have commonly been utilized in past planetary entry trajectory reconstruction, and a new approach for this application that makes use of the pressure measurements. The paper describes assessments of data quality and preprocessing, and results of the flight data reduction from atmospheric entry, which occurred on August 5th, 2012.

Karlgaard, Christopher D.↗
NASA NTRS1980

Transonic unsteady airloads on an energy efficient transport wing with oscillating control surfaces

An aspect-ratio 10.8 supercritical wing with oscillating control surfaces is described. The wing is instrumented with 252 static orifices and 164 in situ dynamic pressure transducers for studying the effects of control surface deflection on steady and unsteady pressures at transonic speeds. Selected results from initial wind-tunnel tests conducted in the Langley Transonic Dynamics Tunnel are discussed. Unsteady pressure results are presented for two trailing-edge control surfaces oscillating separately at the design Mach number of 0.78. Some experimental results are compared with analytical results obtained by using linear lifting-surface theory.

Sandford, M. C.↗
DOE OSTI2022

A Dynamic Pore Network Model for Imbibition Simulation Considering Corner Film Flow

Wetting films can develop in the corners of angular pores under strong wetting conditions. Modeling the dynamics of corner film remains elusive using direct numerical simulations because of the significant scale difference between main meniscus and corner film flow. In this paper, the modified interacting capillary bundle model (ICB), developed in our previous work to describe accurately corner film dynamics in a single square tube, is incorporated into a single-pressure dynamic pore network model (DPNM) to simulate imbibition in strongly wetting porous media with corner film flow. The traditional pore network is decomposed into several layers of interacting subpore networks where the 0th layer of subpore network simulates the main meniscus flow and higher layers the corner film flow. The fluid flow between different layers is captured by interlayer throats. In addition, the snap-off mechanism caused by the thickening of wetting corner film is considered. The accuracy of the developed model is validated for four cases: spontaneous imbibition in a single square tube, wetting fluid redistribution through corner films under a capillary pressure difference, snap off in a narrow throat connecting two large pores, and imbibition dynamics in a real microfluidic porous geometry. The validated model is then used to simulate both spontaneous and controlled imbibition in a pore network with random pore size distribution. Finally, the interaction between corner film and main meniscus flow in porous media is analyzed from a pore-scale perspective.

58 GEOSCIENCES↗
NASA NTRS1985

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfroils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗
NASA NTRS1986

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfols were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a clasical usnteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗
NASA NTRS1990

External nozzle flap dynamic load measurements on F-15 S/MTD model

Dynamic pressure loads were obtained on 1/12 scale models of the F-15B production aircraft and the F-15 S/MTD experimental aircraft with rectangular nozzles and canards. Flight Mach numbers from 0.51 to 1.20 were studied for aircraft angles of attack from 0 to 10 deg and nozzle pressure ratios from 1.00 to 5.09. The results show that dynamic levels are lower in the internozzle region of twin rectangular nozzles than are levels found with twin axisymmetric nozzles. At other locations, the levels associated with both geometries are of the same order of magnitude when normalized by aircraft dynamic Q. At Mach number of 0.51, the loads spectrum is dominated by plume shock noise processes for both geometries. Above Mach 0.51, this mechanism is associated with either vortex bursting from a forward location or turbulent boundary layer separation over the nozzle external flaps. At supersonic speeds both geometries show significantly decreased load levels.

Seiner, John M.↗
DOE OSTI2024

Benchmarking the Suitability of Novec$^{\mathrm{TM}}$ 4710 for Application in Flux Compression Generators

Here, an experimental study evaluated the feasibility of replacing traditional insulating gases such as SF 6 with C 4 F 7 N (3M TM , Novec 4710) in flux compression generator (FCG) applications. Currently available data indicate that Novec 4710 could offer certain performance benefits over SF 6 . However, the available literature is focused on low frequency (50–60 Hz) and dc at static pressures. To evaluate the performance of Novec 4710 under the pulsed dynamic pressure and temperature conditions found in an FCG, we report a performance comparison between three sets of identical FCGs using air, SF 6 , and Novec 4710 as the insulating gas. The generators used in this study had a single stage, directly seeded design with an armature diameter of 25 mm and a stator diameter of 46 mm. To highlight the performance of the different gases rather than any wire insulation, the stator was constructed with uninsulated wire. Furthermore, the generators were seeded aggressively, making the performance difference between the different gases more apparent. The performance was monitored with a pair of differential Rogowski coils that captured the generators’ di / dt while also using high-speed videography to capture possible gaseous breakdown signatures. The data gathered during this study indicate that Novec 4710 performs at least as well as SF 6 in FCG applications, if not significantly better.

42 ENGINEERING↗
NASA NTRS1983

Measuring Small Leak Holes

Hole sizes deduced from pressure measurements. Measuring apparatus consists of pitot tube attached to water-filled manometer. Compartment tested is pressurized with air. Pitot probe placed at known distance from leak. Dynamic pressure of jet measured at that point and static pressure measured in compartment. Useful in situations in which small leaks are tolerable but large leaks are not.

Koch, D. E.↗
NASA NTRS1988

Aerothermal tests of quilted dome models on a flat plate at a Mach number of 6.5

Aerothermal tests were conducted in the NASA Langley 8 Foot High Temperature Tunnel (8'HTT) at a Mach number of 6.5 on simulated arrays of thermally bowed metallic thermal protection system (TPS) tiles at an angle of attack of 5 deg. Detailed surface pressures and heating rates were obtained for arrays aligned with the flow and skewed 45 deg diagonally to the flow with nominal bowed heights of 0.1, 0.2, and 0.4 inch submerged in both laminar and turbulent boundary layers. Aerothermal tests were made at a nominal total temperature of 3300 R, a total pressure of 400 psia, a total enthalpy of 950 Btu/lbm, a dynamic pressure of 2.7 psi, and a unit Reynolds number of 400,000 per foot. The experimental results form a data base that can be used to help protect aerothermal load increases from bowed arrays of TPS tiles.

Glass, Christopher E.↗
NASA NTRS2001

Panel Flutter and Sonic Fatigue Analysis for RLV

A methodology is presented for the flutter analysis of the seal of thermal protection system (TPS) panel of X-33 Advanced Technology Demonstrator test vehicle. The seal is simulated as a two-dimensional cantilevered panel with an elastic stopper, which is modeled as an equivalent spring. This cantilever beam-spring model under the aerodynamic pressure at supersonic speeds turns out to be an impact nonlinear dynamic system. The flutter analysis of the seal is thus carried out using, time domain numerical simulation with a displacement stability criterion. The flutter boundary of the seal is further verified with a family of three traditional and one nontraditional panel flutter models. The frequency domain method that applies eigenanalysis on the traditional panel flutter problem was used. The results showed that the critical dynamic pressure could be more than doubled with properly chosen material for the base stopper. The proposed methodology can be easily extended to three-dimensional panel seals with flow angularity.

Mei, Chuh↗
NASA NTRS2004

Low-Pressure Capability of NASA Glenn's 10- by 10-Foot Supersonic Wind Tunnel Expanded

Extremely low dynamic pressure Q conditions are desired for space-related research including the testing of parachute designs and other decelerator concepts for future vehicles landing on Mars. Therefore, the low-pressure operating capability of the Abe Silverstein 10- by 10-foot Supersonic Wind Tunnel (10 10 SWT) at NASA Glenn Research Center was recently increased. Successful checkout tests performed in the fall of 2002 showed significantly reduced minimum operating pressures in the wind tunnel.

Roeder, James W.↗