Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “STAGNATION 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 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Turbomachinery for Low-to-High Mach Number Flight

The thrust capability of turbojet cycles is reduced at high flight Mach number (3+) by the increase in inlet stagnation temperature. The 'hot section' temperature limit imposed by materials technology sets the maximum heat addition and, hence, sets the maximum flight Mach number of the operating envelope. Compressor pre-cooling, either via a heat exchanger or mass-injection, has been suggested as a means to reduce compressor inlet temperature and increase mass flow capability, thereby increasing thrust. To date, however, no research has looked at compressor cooling (i.e., using a compressor both to perform work on the gas path air and extract heat from it simultaneously). We wish to assess the feasibility of this novel concept for use in low-to-high Mach number flight. The results to-date show that an axial compressor with cooling: (1) relieves choking in rear stages (hence opening up operability), (2) yields higher-pressure ratio and (3) yields higher efficiency for a given corrected speed and mass flow. The performance benefit is driven: (i) at the blade passage level, by a decrease in the total pressure reduction coefficient and an increase in the flow turning; and (ii) by the reduction in temperature that results in less work required for a given pressure ratio. The latter is a thermodynamic effect. As an example, calculations were performed for an eight-stage compressor with an adiabatic design pressure ratio of 5. By defining non-dimensional cooling as the percentage of compressor inlet stagnation enthalpy removed by a heat sink, the model shows that a non-dimensional cooling of percent in each blade row of the first two stages can increase the compressor pressure ratio by as much as 10-20 percent. Maximum corrected mass flow at a given corrected speed may increase by as much as 5 percent. In addition, efficiency may increase by as much as 5 points. A framework for characterizing and generating the performance map for a cooled compressor has been developed. The approach is based upon CFD computations and mean line analysis. Figures of merit that characterize the bulk performance of blade passage flows with and without cooling are extracted from CFD solutions. Such performance characterization is then applied to a preliminary compressor design framework (mean line). The generic nature of this approach makes it suitable for assessing the effect of different types of compressor cooling schemes, such as heat exchange or evaporative cooling (mass injection). Future work will focus on answering system level questions regarding the feasibility of compressor cooling. Specifically, we wish to determine the operational parametric space in which compressor cooling would be advantageous over other high flight Mach number propulsion concepts. In addition, we will explore the design requirements of cooled compressor turbomachinery, as well as the flow phenomena that limit and control its operation, and the technology barriers that must be crossed for its implementation.

Tan, Choon S.↗
NASA NTRS2012

Investigation of Hypersonic Laminar Heating Augmentation in the Stagnation Region

Laminar stagnation region heating augmentation is investigated in the AEDC Tunnel 9 at Mach 10 by performing high frequency surface pressure and heat transfer measurements on the Orion CEV capsule at zero degree angle-of-attack for unit Reynolds numbers between 0.5 and 15 million per foot. Heating augmentation increases with Reynolds number, but is also model size dependent as it is absent on a 1.25-inch diameter model at Reynolds numbers where it reaches up to 15% on a 7-inch model. Heat transfer space-time correlations on the 7-inch model show that disturbances convect at the boundary layer edge velocity and that the streamwise integral scale increases with distance. Therefore, vorticity amplification due to stretching and piling-up in the stagnation region appears to be responsible for the stagnation point heating augmentation on the larger model. This assumption is reinforced by the f(exp -11/3) dependence of the surface pressure spectrum compared to the f(exp -1) dependence in the free stream. Vorticity amplification does not occur on the 1.25- inch model because the disturbances are too large. Improved free stream fluctuation measurements will be required to determine if significant vorticity is present upstream or mostly generated behind the bow shock.

Marineau, Eric C.↗
NASA NTRS1964

Injection and distribution of liquids in the flow fields of blunt shapes at hypersonic speeds

A possible method for reducing the attenuation of radio transmission from hypersonic vehicles is by injection of a liquid into the flow field. The hot plasma layer surrounding the vehicle could presumably be quenched if the latent heat of the liquid and the heat capacity of the vapor could be utilized efficiently. In the case of stagnation-point injection, the momentum of the liquid could be utilized to modify the bow shock and possibly reduce the initial ionization. To attain these objectives of efficient coolant utilization or modification of the bow shock a knowledge of the liquid injection, atomization, spray penetration, and evaporation processes is required. This report presents the results of a study of liquid injection into hypersonic flow fields. A simplified theory for evaporation distances is derived and is shown to be in qualitative agreement with experimental data. Correlation of experimental data for maximum cross-current penetration of the liquid spray for injection near the shoulder pf a hemisphere cylinder, a sphere 9° cone, and a Mercury model are given. These data were obtained in the Langley Mach 8 variable density tunnel with both liquid nitrogen and water injection to provide a large range of liquid parameters. Flow-field disturbances, as indicated by schlieren photographs and surface pressures, increased with the ratio of coolant mass flux to airstream mass flux. For stagnation-point injection, the maximum shock standoff distance was correlated by the ratio of the liquid jet momentum flux to the airstream momentum flux.

Ivan E. Beckwith↗
NASA NTRS1975

Analysis of atmospheric flow over a surface protrusion using the turbulence kinetic energy equation

Atmospheric flow fields resulting from a semi-elliptical surface obstruction in an otherwise horizontally homogeneous statistically stationary flow are modelled with the boundary-layer/Boussinesq-approximation of the governing equation of fluid mechanics. The turbulence kinetic energy equation is used to determine the dissipative effects of turbulent shear on the mean flow. Mean-flow results are compared with those given in a previous paper where the same problem was attacked using a Prandtl mixing-length hypothesis. Iso-lines of turbulence kinetic energy and turbulence intensity are plotted in the plane of the flow. They highlight regions of high turbulence intensity in the stagnation zone and sharp gradients in intensity along the transition from adverse to favourable pressure gradient.

Frost, W.↗
NASA NTRS1975

Analysis of atmospheric flow over a surface protrusion using the turbulence kinetic energy equation with reference to aeronautical operating systems

Flow over surface obstructions can produce significantly large wind shears such that adverse flying conditions can occur for aeronautical systems (helicopters, STOL vehicles, etc.). Atmospheric flow fields resulting from a semi-elliptical surface obstruction in an otherwise horizontally homogeneous statistically stationary flow are modelled with the boundary-layer/Boussinesq-approximation of the governing equation of fluid mechanics. The turbulence kinetic energy equation is used to determine the dissipative effects of turbulent shear on the mean flow. Iso-lines of turbulence kinetic energy and turbulence intensity are plotted in the plane of the flow and highlight regions of high turbulence intensity in the stagnation zone and sharp gradients in intensity along the transition from adverse to favourable pressure gradient. Discussion of the effects of the disturbed wind field in CTOL and STOL aircraft flight path and obstruction clearance standards is given. The results indicate that closer inspection of these presently recommended standards as influenced by wind over irregular terrains is required.

Frost, W.↗
NASA NTRS1977

Two phase choke flow in tubes with very large L/D

Data were obtained for two phase and gaseous choked flow nitrogen in a long constant area duct of 16200 L/D with a diverging diffuser attached to the exit. Flow rate data were taken along five isotherms (reduced temperature of 0.81, 0.96, 1.06, 1.12, and 2.34) for reduced pressures to 3. The flow rate data were mapped in the usual manner using stagnation conditions at the inlet mixing chamber upstream of the entrance length. The results are predictable by a two phase homogeneous equilibrium choking flow model which includes wall friction. A simplified theory which in essence decouples the long tube region from the high acceleration choking region also appears to predict the data resonably well, but about 15 percent low.

Hendricks, R. C.↗
NASA NTRS1977

Two phase choke flow in tubes with very large L/D

Two phase and gaseous choked flow data for fluid nitrogen were obtained for a test section which was a long constant area duct of 16 200 L/D with a diverging diffuser attached to the exit. Flow rate data were taken along five isotherms (reduced temperature of 0.81, 0.96, 1.06, 1.12, and 2.34) for reduced pressures to 3. The flow rate data were mapped in the usual manner using stagnation conditions at the inlet mixing chamber upstream of the entrance length. The results are predictable by a two-phase homogeneous equilibrium choking flow model which includes wall fraction. A simplified theory which in essence decouples the long tube region from the high acceleration choking region also appears to predict the data reasonably well, but about 15 percent low.

Hendricks, R. C.↗
NASA NTRS1980

Shock-tube studies of radiative base heating of Jovian probe

A 6.4-cm-diameter scale model of the Jovian entry vehicle is tested in an electric-arc-driven shock tube and a 5-cm-diameter sphere model is tested in a combustion-driven shock tube and in an electric-arc-driven shock tunnel. The radiative heat-transfer rate and pressure on the front and the base regions are measured in the absence of ablation with sensors imbedded in the models in a stream consisting of 10% hydrogen in a bath of either neon or argon. The measured radiative heat-transfer rates and pressures range to about 22 kW/sq cm and 12 atm, respectively, at the front stagnation point. The ratio of the radiative heat-transfer rate at the base stagnation point to that at the front stagnation point is found to be about 1/4 for the sphere at Mach 1.8, about 1/30 for the sphere at Mach 4.8, and about 1/6 for the scale model at Mach 1.7. The present experimental results agree well with the theoretical predictions of Park, thus indicating that Park's theory is valid.

Shirai, H.↗
NASA NTRS1994

Pressure distributions on a 0.02-scale Space Shuttle orbiter nose at Mach 21.5 in helium

Pressure distributions on a 0.02-scale model of the Space Shuttle orbiter forward fuselage were obtained in the 22-inch aerodynamic leg of the Langley Hypersonic Helium Tunnel Facility at a nominal free-stream Mach number of 21.5 and a ratio of specific heats of 1.67 for inclusion in the database of the Shuttle entry air data system (SEADS). The data were measured at model angles of attack of 0 deg to 50 deg in 5 deg increments for zero sideslip angle and at model sideslip angles of -5 deg to 5 deg for angles of attack equal to 5, 20, 35, and 40 deg. These data displayed trends similar to those observed in other wind tunnels at Mach 6 and 10 in air. Specifically noted is a shift in the location of the stagnation point at angles of attack above 15 deg; this effect did not, however, occur in flight. By comparison, the data obtained at Mach 6 in the Langley Hypersonic CF4 Tunnel, corresponding to a lower ratio of specific heats in the postshock region than those in helium and air, showed some reduction of the stagnation point shift at the higher angles of attack. The differences between flight and wind tunnel pressure distributions are believed due primarily to high-temperature gas chemistry effects in flight, which include lower effective specific heat ratios but which were not completely duplicated in the wind tunnels.

Ashby, George C., Jr.↗
NASA NTRS1971

Thermodynamic properties and theoretical rocket performance of hydrogen to 100,000 K and 1.01325 x 10 to the 8th power N/sq m

The composition and thermodynamic properties were calculated for 100 to 110,000 K and 1.01325 x 10 to the 2nd power to 1.01325 x 10 to the 8th power N/sq m for chemical equilibrium in the Debye-Huckel and ideal-gas approximations. Quantities obtained were the concentrations of hydrogen atoms, protons, free electrons, hydrogen molecules, negative hydrogen ions, hydrogen diatomic molecular ions, and hydrogen triatomic molecular ions, and the enthalpy, entropy, average molecular weight, specific heat at constant pressure, density, and isentropic exponent. Electronically excited states of H and H2 were included. Choked, isentropic, one-dimensional nozzle flow with shifting chemical equilibrium was calculated to the Debye-Huckel and ideal-gas approximations for stagnation temperatures from 2500 to 100,000 K. The mass flow per unit throat area and the sonic flow factor were obtained. The pressure ratio, temperature, velocity, and ideal and vacuum specific impulses at the throat and for pressure ratios as low as 0.000001 downstream were found. For high temperatures at pressures approaching 1.01325 x 10 to the 8th power N/sq m, the ideal-gas approximation was found to be inadequate for calculations of composition, precise thermodynamic properties, and precise nozzle flow. The greatest discrepancy in nozzle flow occurred in the exit temperature, which was as much as 21 percent higher when the Debye-Huckel approximation was used.

Patch, R. W.↗
NASA NTRS1991

Multiple shock-shock interference on a cylindrical leading edge

The details of an experimental study of shock wave interference heating on a cylindrical leading edge representative of the cowl of a rectangular hypersonic engine inlet are presented. This Mach 8 study has provided the first detailed pressure and heat transfer rate distributions on a cylinder resulting from a two-dimensional shockwave interference pattern created by two incident oblique shock waves intersecting the cylinder bow shock wave. The peak heat transfer rate was 38 times the undisturbed flow stagnation point level and occurred when the two oblique shock waves coalesced prior to intersecting the cylinder bow shock wave. Development of pressure deflection diagrams identified a new interference pattern consisting of concomitant supersonic jets separated from each other by a shear layer and submerged in the subsonic region between the bow shock wave and body.

Wieting, Allan R.↗
NASA NTRS1995

Strut and wall interference on jet-induced ground effects of a STOVL aircraft in hover

A small scale ground effect test rig was used to study the ground plane flow field generated by a STOVL aircraft in hover. The objective of the research was to support NASA-Ames Research Center planning for the Large Scale Powered Model (LSPM) test for the ARPA-sponsored ASTOVL program. Specifically, small scale oil flow visualization studies were conducted to make a relative assessment of the aerodynamic interference of a proposed strut configuration and a wall configuration on the ground plane stagnation line. A simplified flat plate model representative of a generic jet-powered STOVL aircraft was used to simulate the LSPM. Cold air jets were used to simulate both the lift fan and the twin rear engines. Nozzle Pressure Ratios were used that closely represented those used on the LSPM tests. The flow visualization data clearly identified a shift in the stagnation line location for both the strut and the wall configuration. Considering the experimental uncertainty, it was concluded that either the strut configuration o r the wall configuration caused only a minor aerodynamic interference.

Kristy, Michael H.↗
NASA NTRS1979

Experimental studies of radiative base heating of a Jovian entry model

A scale model of the Jovian entry vehicle of 6.4 cm diameter is tested in an electric-arc-driven shock tube while a sphere model of 5 cm diameter is tested in a combustion-driven shock tube and an electric-arc-driven shock tunnel. The radiative heat-transfer rate and pressure on the front and the base regions are measured in the absence of ablation with sensors imbedded in the models in a stream consisting of 10% hydrogen in a bath of either neon or argon. The measured radiative heat-transfer rates and pressures are up to about 22 kW/sq cm and 12 atm, respectively, at the front stagnation point. The ratio of the base-to-stagnation-point radiative heat-transfer rates is found to be about 1/4 for the sphere at Mach 1.8, about 1/30 for the sphere at Mach 4.8, and about 1/6 for the scale model at Mach 1.7. When the effects of model geometry and Mach number are accounted for, the present experimental results agree well with the theoretical prediction of Park.

Shirai, H.↗
NASA NTRS1988

Convection and mass-transport in laser-induced chemical vapor deposition

Gas flow and energy and species transport in laser-induced chemical vapor deposition (LICVD) of amorphous silicon films by silane pyrolysis are analyzed by finite element analysis of a two-dimensional model for the process. Spatial nonuniformity of the deposited film is shown to result from diffusion controlled transport of products between the beam and substrate. Deposition profiles are affected by buoyancy-driven convection only at increased gas pressures. Horizontal orientation of the reactor with respect to gravity is optimal because the stagnation-like flow, that results adjacent to the substrate, enhances mixing, and smoothes the film profile.

Patnaik, S.↗
NASA NTRS1988

Progress towards the development of an inviscid-viscous interaction method for unsteady flows in turbomachinery cascades

The velocity and heat transfer characteristics of blade passage flows subject to onset velocities that vary in space and time are calculated by means of a general method involving the solution of inviscid and boundary-layer equations coupled with the Hilbert integral. The performance of this boundary layer method for flows with prescribed pressure distribution is evaluated in terms of two model problems that give attention to the stagnation region. One model corresponds to an oscillating airfoil with moving stagnation point; the other is a simulation of a blade which is subject to a uniform onset velocity.

Cebeci, T.↗
NASA NTRS2003

Film Cooling Flow Effects on Post-Combustor Trace Chemistry

Film cooling injection is widely applied in the thermal design of turbomachinery, as it contributes to achieve higher operating temperature conditions of modern gas turbines, and to meet the requirements for reliability and life cycles. It is a significant part of the high-pressure turbine system. The film cooling injection, however, interacts with the main flow and is susceptible to have an influence on the aerodynamic performance of the cooled components, and through that may cause a penalty on the overall efficiency of the gas turbine. The main reasons are the loss of total pressure resulting from mixing the cooling air with mainstream and the reduction of the gas stagnation temperature at the exit of the combustion chamber to a lower value at the exit of nozzle guide vane. In addition, the impact of the injected air on the evolution of the trace species of the hot gas is not yet quite clear. This work computationally investigates the film cooling influence on post-combustor trace chemistry, as trace species in aircraft exhaust affect climate and ozone.

Wey, Thomas↗
NASA NTRSDate not supplied

CFD Simulations of the IHF 13-Inch Nozzle Flow: 55-deg Sphere-Cone Model, Manufactured Fences and Gaps

This paper reports computational analyses of testsin a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using 55-deg sphere-cone models placed in a free jet downstream of the 13-inch diameter conical nozzle in the Ames 60-MW Interaction Heating Facility. Some of the sphere-cone models include surface features such as manufactured fences and gaps intended to simulate effects of differential recession, all of which disturb the flow, producing augmented heating locally and downstream.Test calibration data were obtained using slug and Gardon gage stagnation calorimeters, and a sphere-cone calorimeter model with six Gardon gages and five pressure tabs.The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, including the models tested, and comparisons with the experimental measurements. These simulations take into account nonuniform total enthalpy and mass flux profiles at the nozzle inlet as well as the expansion waves emanating from the nozzle exit and their effects on the model flowfields.

Arc-jets↗
NASA NTRSDate not supplied

CFD Simulations of the IHF 13-Inch Nozzle Flow: 55-deg Sphere-Cone Model, Manufactured Fences and Gaps

This paper reports computational analyses of tests in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using 55-deg sphere-cone models placed in a free jet downstream of the 13-inch diameter conical nozzle in the Ames 60-MW Interaction Heating Facility. Some of the sphere-cone models include surface features such as manufactured fences and gaps intended to simulate effects of differential recession, all of which disturb the flow, producing augmented heating locally and downstream. Test calibration data were obtained using slug and Gardon gage stagnation calorimeters, and a sphere-cone calorimeter model with six Gardon gages and five pressure tabs. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, including the models tested, and comparisons with the experimental measurements. These simulations take into account nonuniform total enthalpy and mass flux profiles at the nozzle inlet as well as the expansion waves emanating from the nozzle exit and their effects on the model flowfields.

Arc-jets↗