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At least 235 records · Page 13

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NASA NTRS1988

Cryogenic wind tunnel research - A global perspective

The development of cryogenic wind tunnels is reviewed and 13 cryogenic wind tunnels currently operating in England, France, Germany, Japan, and the U.S. are described. A table illustrating the characteristics of these tunnels is presented, including test gases, test section sizes, speed ranges, stagnation pressure and temperature, and running time. The research conducted using the various wind tunnels is outlined and the operation of each of the tunnels is considered.

Dress, D. A.↗

NASA NTRS1989

Mechanistic model for catalytic recombination during aerobraking maneuvers

Several mechanistic models are developed to predict recombination coefficients for use in heat shield design for reusable surface insulation (RSI) on aerobraking vehicles such as space shuttles. The models are applied over a temperature range of 300 to 1800 K and a stagnation pressure range of 0 to 3,000 Pa. A four parameter model in temperature was found to work best; however, several models (including those with atom concentrations at the surface) were also investigated. Mechanistic models developed with atom concentration terms may be applicable when sufficient data becomes available. The requirement is shown for recombination experiments in the 300 to 1000 K and 1500 to 1850 K temperature range, with deliberate concentration variations.

Willey, Ronald J.↗

NASA NTRS1987

Cryogenic wind tunnels. II

The application of the cryogenic concept to various types of tunnels including Ludwieg tube tunnel, Evans clean tunnel, blowdown, induced-flow, and continuous-flow fan-driven tunnels is discussed. Benefits related to construction and operating costs are covered, along with benefits related to new testing capabilities. It is noted that cooling the test gas to very low temperatures increases Reynolds number by more than a factor of seven. From the energy standpoint, ambient-temperature fan-driven closed-return tunnels are considered to be the most efficient type of tunnel, while a large reduction in the required tunnel stagnation pressure can be achieved through cryogenic operation. Operating envelopes for three modes of operation for a cryogenic transonic pressure tunnel with a 2.5 by 2.5 test section are outlined. A computer program for calculating flow parameters and power requirements for wind tunnels with operating temperatures from saturation to above ambient is highlighted.

Kilgore, Robert A.↗

NASA NTRS1991

Condensation effects on Rayleigh scattering measurements in a supersonic wind tunnel

Rayleigh and Raman scattering measurements have been performed in the 20-in., Mach 6 wind tunnel at the NASA Langley Research Center. Rayleigh results show signal levels much higher than expected for molecular scattering in the tunnel, whereas densities deduced from spontaneous Raman scattering of molecular nitrogen are in good agreement with the expected nitrogen densities in the facility. The apparent discrepancy in the Rayleigh result is attributed to cluster formation as a result of expansion in the tunnel. The dependence of the Rayleigh signal on the stagnation pressure and temperature is also discussed.

Shirinzadeh, B.↗

NASA NTRS1991

Phenomenological and experimental study of the thermal response of low density silica ablators to high enthalpy plasma flows

A complete analysis of the melting-ablator governing equations is carried out for porous ablators with the purpose of establishing an approximate analytical model capable of determining surface recession, temperature, and heat flux as well as in-depth temperature profiles. A detailed dimensional analysis of the melt-flow governing equations is performed for parameters applicable to high-enthalpy flows. Results of arc-jet flow experiments performed on NASA Reusable Surface Insulation materials of different densities and over a range of stagnation pressures and heat fluxes are discussed. A dimensionless correlation for melt run-offs is derived in terms of material and flow parameters; however, it is pointed out that the accuracy of the current data is sufficient to conclude the compliance with the model.

Henline, William D.↗

NASA NTRS1991

The magnetic barrier at Venus

Altitude profiles of the Venus magnetic barrier are derived here from a statistical analysis of the Pioneer Venus Orbiter magnetometer data. The outer boundary of the magnetic barrier is then compared with the obstacle expected from gasdynamic models of the bow shock, and the stagnation pressure is compared with that expected from gasdynamic theory. The magnetic barrier is strongest at the subsolar point and weakens as expected with increasing solar zenith angle. The existence of a north-south asymmetry in the barrier strength is also demonstrated. The magnetic barrier is about 200 km thick at the subsolar point and 800 km thick at the terminator. The magnetic barrier transfers most of the solar wind dynamic pressure to the ionosphere via the enhanced magnetic pressure. The convected field gasdynamic model predicts the correct bow shock location if the magnetic barrier is treated as the obstacle.

Zhang, T. L.↗

NASA NTRS1992

Investigation of three-dimensional flow field in a turbine including rotor/stator interaction. II - Three-dimensional flow field at the exit of the nozzle

The nozzle exit flowfield was measured at two axial locations with a miniature five-hole probe. Measurements were taken from hub-to-tip, blade-to-blade at 21 radial locations and at two axial locations downstream of the nozzle trailing edge to resolve the flowfield accurately including the nozzle wake, secondary flow region, horseshoe vortex and losses. All three components of the velocity, stagnation pressure, static pressure, and pitch and yaw angles have been resolved very accurately. The wake data seems to indicate that the decay of the wake is faster than the wake of an isolated nozzle row. The cause of this is attributed to the presence of the rotor downstream. A distinct vortex core has been observed near the tip. The indications are that the horseshoe vortex and the passage vortex have merged to produce a single loss core region. Roughly a third of the blade height passage near the tip and a third of the blade height near the hub is dominated by secondary flow, passage vortex and the horseshoe vortex phenomena. Only the middle third of the nozzle behaves as per design. These and other data are presented, interpreted and synthesized to understand the nozzle flowfield.

Zaccaria, M.↗

NASA NTRS1993

Development and operation of new arc heater technology for a large-scale scramjet propulsion test facility

An arc-heater technology development effort conducted at NASA-Ames has led to the creation of a 100 MW Huels arc heater which has been integrated into the hydrogen-fueled Direct Connect Arcjet Facility for long-duration, high Mach-number scramjet performance evaluation. This development effort has significantly advanced the design of vortex-stabilized arc heaters; air enthalpy levels of 2.8-7.5 MJ/kg, at stagnation pressures of up to 45 atm, are produced. The facility furnishes technology-development support that is critical to the definition of NASP-related propulsion systems.

Balboni, John↗

NASA NTRS1993

Atmospheric entry of Mars-return nuclear-powered vehicles due to accidental termination of operations

The entry of nuclear reactors into Earth's atmosphere resulting from an accidental or inadvertent abort of a space vehicle powered by nuclear-thermal rockets is investigated. The study is made for a typical piloted Mars mission vehicle incapacitated by an accident or malfunction during the Earth-arrival phase of the Mars-return journey due to simultaneous, multiple failures of its component systems. A single accident/abort scenario resulting in three entry possibilities is considered for a nominal hyperbolic in-bound approach velocity of 8 km/sec. The most severe case involving a direct entry is then analyzed over a broad range of approach velocities extending to 12 km/sec to include sprint-type missions. The results indicate that the severe surface heating, stagnation pressures, and g-loads are greater than 150 kW/sq cm, 300 atm, and 800-g, respectively. The wall heat transfer rate exceeds the value that can be accommodated by a carbon heatshield through radiation equilibrium prior to sublimation at 5500 K. These conditions are beyond our previous experience in crew safety, structural design, and thermal protection.

Menees, Gene P.↗

NASA NTRS1951

Experimental investigation of an 0.8 hub-tip radius-ratio,nontwisted-rotor-blade turbine

An experimental investigation of a 0.8 hub-tip radius ratio, nontwisted-rotor-blade turbine designed for a stagnation-pressure ratio of 2.5 and an equivalent mean blade speed of 643 feet per second was made in a cold-air turbine with (a) nontwisted stator blades, and (b) twisted stator blades designed to maintain zero rotor-inlet incidence angles. Turbine efficiencies of the order of 0.85 at the design point were obtained with a nontwisted-rotor-blade turbine with a hub-tip radius ratio of 0.80. The turbine with the twisted stator blades gave higher efficiencies at the design point (of the order of 1.5 percentage points) than the turbine with the nontwisted stator blades.

TURBINES - AXIAL FLOW↗

NASA NTRS1949

Criterions for condensation-free flow in supersonic tunnels

The results of an investigation of water-vapor condensation shocks in the air passing through supersonic tunnels are presented. Criterions for condensation-free flow are established by correlating experimental observations with the Volmer theory of nuclei formation. Experimental observations were made at Mach numbers up to 2.01. The criterions are presented in a form independent of tunnel-inlet stagnation pressure and are extended theoretically to a Mach number of 4.00. Preliminary evidence of the effect of tunnel size on the criterion is presented.

Burgess, Warren C↗

NASA NTRS1993

Termination shock - Solar cycle variations of location and speed

Plasma and magnetic field data from Voyagers 1 and 2 over an 11-year period (1978-1988) together with a one-fluid MHD model are used to study the motion of the termination shock. The observed large fluctuations in solar wind parameters cause variations in the speed, in the jump conditions, and in the location of the termination shock. If the interstellar stagnation pressure is equivalent to a magnetic field of 0.5 nT plus draping enhancements, the termination shock moves between 88 AU and 102 AU. The location of the termination shock is anticorrelated with the sunspot number. The termination shock is very strong with pressure ratios of a few thousands, the fast Mach number is about 12, and the plasma beta ratio jumps from about 0.1 to about 14. During the declining phase of the solar cycle, the shock speed is often greater than or comparable to the spacecraft speed. If the crossing of the termination shock by a spacecraft occurs during the declining phase, the spacecraft will cross the termination shock 3 or more times over a period of about 4 years.

Whang, Y. C.↗

NASA NTRS1994

Design of a variable area diffuser for a 15-inch Mach 6 open-jet tunnel

The Langley 15-inch Mach 6 High Temperature Tunnel was recently converted from a Mach 10 Hypersonic Flow Apparatus. This conversion was effected to improve the capability of testing in Mach 6 air at relatively high reservoir temperatures not previously possible at Langley. Elevated temperatures allow the matching of the Mach numbers, Reynolds numbers, and ratio of wall-to-adiabatic-wall temperatures (TW/Taw) between this and the Langley 20-inch Mach 6 CF4 Tunnel. This ratio is also matched for Langley's 31-inch Mach 10 Tunnel and is an important parameter useful in the simulation of slender bodies such as National Aerospace Plane (NASP) configurations currently being studied. Having established the nozzle's operating characteristics, the decision was made to install another test section to provide model injection capability. This test section is an open-jet type, with an injection system capable of injecting a model from retracted position to nozzle centerline between 0.5 and 2 seconds. Preliminary calibrations with the new test section resulted in Tunnel blockage. This blockage phenomenon was eliminated when the conical center body in the diffuser was replaced. The issue then, is to provide a new and more efficient variable area diffuser configuration with the capability to withstand testing of larger models without sending the Tunnel into an unstart condition. Use of the 1-dimensional steady flow equation with due regard to friction and heat transfer was employed to estimate the required area ratios (exit area / throat area) in a variable area diffuser. Correlations between diffuser exit Mach number and area ratios, relative to the stagnation pressure ratios and diffuser inlet Mach number were derived. From these correlations, one can set upper and lower operating pressures and temperatures for a given diffuser throat area. In addition, they will provide appropriate input conditions for the full 3-dimensional computational fluid dynamics (CFD) code for further simulation studies.

Loney, Norman W.↗

NASA NTRS1995

Leading edge film cooling effects on turbine blade heat transfer

An existing three dimensional Navier-Stokes code, modified to include film cooling considerations, has been used to study the effect of spanwise pitch of shower-head holes and coolant to mainstream mass flow ratio on the adiabatic effectiveness and heat transfer coefficient on a film-cooled turbine vane. The mainstream is akin to that under real engine conditions with stagnation temperature = 1900 K and stagnation pressure = 3 MPa. It is found that with the coolant to mainstream mass flow ratio fixed, reducing P, the spanwise pitch for shower-head holes, from 7.5 d to 3.0 d, where d is the hole diameter, increases the average effectiveness considerably over the blade surface. However, when P/d= 7.5, increasing the coolant mass flow increases the effectiveness on the pressure surface but reduces it on the suction surface due to coolant jet lift-off. For P/d = 4.5 or 3.0, such an anomaly does not occur within the range of coolant to mainstream mass flow ratios analyzed. In all cases, adiabatic effectiveness and heat transfer coefficient are highly three-dimensional.

Garg, Vijay K.↗

NASA NTRS1995

A Numerical Study of the Effect of Wake Passing on Turbine Blade Film Cooling

Time-accurate and steady three-dimensional viscous turbulent numerical simulations were performed to study the effect of upstream blade wake passing unsteadiness on the performance of film cooling on a downstream axial turbine blade. The simulations modeled the blade as spanwise periodic and of infinite span. Both aerodynamic and heat transfer quantities were explored. A showerhead film cooling arrangement typical of modern gas turbine engines was employed. Showerhead cooling was studied because of its anticipated strong sensitivity to upstream flow fluctuations. The wake was modeled as a region of zero axial velocity on the upstream computational boundary which translated with each iteration. This model is compatible with a planned companion experiment in which the wakes will be produced by a rotating row of cylindrical rods upstream of an annular turbine cascade. It was determined that a steady solution with appropriate upstream swirl and stagnation pressure predicted the span-average film effectiveness quite well. The major difference is a 2 to 3 percent overprediction of span-average film effectiveness by the steady simulation on the pressure surface and in the showerhead region. Local overpredictions of up to 8 percent were observed in the showerhead region. These differences can be explained by the periodic relative lifting of the boundary layer and enhanced mixing in the unsteady simulations.

Heidmann, James D.↗

NASA NTRS1997

Effect of Coolant Temperature and Mass Flow on Film Cooling of Turbine Blades

A three-dimensional Navier Stokes code has been used to study the effect of coolant temperature, and coolant to mainstream mass flow ratio on the adiabatic effectiveness of a film-cooled turbine blade. The blade chosen is the VKI rotor with six rows of cooling holes including three rows on the shower head. The mainstream is akin to that under real engine conditions with stagnation temperature = 1900 K and stagnation pressure = 3 MPa. Generally, the adiabatic effectiveness is lower for a higher coolant temperature due to nonlinear effects via the compressibility of air. However, over the suction side of shower-head holes, the effectiveness is higher for a higher coolant temperature than that for a lower coolant temperature when the coolant to mainstream mass flow ratio is 5% or more. For a fixed coolant temperature, the effectiveness passes through a minima on the suction side of shower-head holes as the coolant to mainstream mass flow, ratio increases, while on the pressure side of shower-head holes, the effectiveness decreases with increase in coolant mass flow due to coolant jet lift-off. In all cases, the adiabatic effectiveness is highly three-dimensional.

Garg, Vijay K.↗

NASA NTRS1997

Modeling of Water Injection into a Vacuum

A loosely coupled two-phase vacuum water plume model has been developed. This model consists of a source flow model to describe the expansion of water vapor, and the Lagrangian equations of motion for particle trajectories. Gas/Particle interaction is modeled through the drag force induced by the relative velocities. Particles are assumed traveling along streamlines. The equations of motion are integrated to obtain particle velocity along the streamline. This model has been used to predict the mass flux in a 5 meter radius hemispherical domain resulting from the burst of a water jet of 1.5 mm in diameter, mass flow rate of 24.2 g/s, and stagnation pressure of 21.0 psia, which is the nominal Orbiter water dump condition. The result is compared with an empirical water plume model deduced from a video image of the STS-29 water dump. To further improve the model, work has begun to numerically simulate the bubble formation and bursting present in a liquid stream injected into a vacuum. The technique of smoothed particle hydrodynamics was used to formulate this simulation. A status and results of the on-going effort are presented and compared to results from the literature.

Alred, John W.↗

NASA NTRS1998

Transmission and Incidence Losses for a Slotted Plate

The objective of this work is to find a model of the stagnation pressure loss resulting from flow through a slotted plate, which is effectively a cascade of flat plate airfoils, particularly at very large angles of incidence. Data from a published experiment is examined, and compared with control volume analysis, and CFD code calculations. An assumption that the loss can be separated into a transmission loss and an incidence loss seems to be justified by the data. Both the data and the CFD code results are consistent with an incidence loss model in which the flow component normal to the slot axis is lost. However, the experimental transmission loss is much larger than calculated values.

Wilson, Jack↗