Engineering topics

Kilgore, R. A.

Publications and source records attributed to Kilgore, R. A..

At least 37 records · Page 2

NASA NTRS1980

Experience in the control of a continuous flow cyrogenic tunnel

The concept of using liquid nitrogen to cool the test gas of a wind tunnel to cryogenic temperatures lead to the achievement of very high Reynolds number flows in relatively small transonic tunnels. The economical operation of liquid nitrogen cooled cryogenic tunnels is critically dependent on fast and accurate control of the tunnel variables. The control problem of a continuous flow fan driven cryogenic tunnel was addressed, first by developing a lumped multivariable mathematical model of a tunnel and validating the model by reconciling the responses of the Langley 0.3 m transonic cryogenic tunnel to the responses of the mathematical model on a simulator. Finally, the development of laws for the closed loop control of the tunnel pressure and temperature and the successful implementation of a control system for the 0.3 m transonic cryogenic tunnel based on these laws are presented.

Kilgore, R. A.↗

NASA NTRS1980

Instrumentation for calibration and control of a continuous-flow cryogenic tunnel

Those aspects of selection and application of calibration and control instrumentation that are influenced by the extremes in the temperature environment to be found in cryogenic tunnels are described with emphasis on the instrumentation and data acquisition system used in the Langley 0.3 m transonic cryogenic tunnel. Typical calibration results obtained in a 20 by 60 cm two dimensional test section are included.

Ladson, C. L.↗

NASA NTRS1979

Development of the cryogenic tunnel concept and application to the US National Transonic Facility

A fan-driven, high Reynolds number, transonic cryogenic wind tunnel is described. The tunnel has a 2.5- by 2.5 m test section and is capable of operating from ambient to cryogenic temperatures at stagnation pressures up to 8.8 atm. An overview of the cryogenic concept and discussions of drive power requirements, support systems, and operating characteristics and performance of the wind tunnel are included.

Kilgore, R. A.↗

NASA NTRS1979

Full-scale aircraft simulation with cryogenic tunnels and status of the National Transonic Facility

The effect of thermal and caloric imperfections in cryogenic nitrogen on boundary layers was determined to indicate that in order to simulate nonadiabatic laminar or turbulent boundary layers in a cryogenic nitrogen wind tunnel, the flight enthalpy ratio, rather than the temperature ratio, should be reproduced. The absence of significant real gas effects on both viscous and inviscid flows makes it unlikely that there will be large real gas effects on the cryogenic tunnel simulation of shock boundary layer interactions or other complex flow conditions encountered in flight. Condensation effects were studied to determine the minimum usable temperature and indicated that under most circumstances free stream Mach number rather than maximum local Mach number determines the onset of condensation effects.

Kilgore, R. A.↗

NASA NTRS1979

Full scale aircraft simulation with cryogenic tunnels and status of the National Transonic Facility

The paper reviews the results of some of the real-gas studies made at Langley that are directly related to establishing the range of operating conditions that can be used in a cryogenic nitrogen wind tunnel and still be assured of valid full-scale simulation. Consideration is given to the important aerodynamic features, anticipated performance capability, status of construction, and projected operating data for the National Transonic Facility.

Kilgore, R. A.↗

NASA NTRS1977

Specific cooling capacity of liquid nitrogen

The assumed cooling process and the method used to calculate the specific cooling capacity of liquid nitrogen are described, and the simple equation fitted to the calculated specific cooling capacity data, together with the graphical form calculated values of the specific cooling capacity of nitrogen for stagnation temperatures from saturation to 350 K and stagnation pressures from 1 to 10 atmospheres, are given.

Kilgore, R. A.↗

NASA NTRS1977

Cryogenic wind-tunnel technology

The cryogenic concept and the advantages it offers with respect to achieving full scale Reynolds number in a moderate size tunnel at reasonable levels of dynamic pressure are described. Aspects which must be considered during the development of a facility that uses gaseous nitrogen as the test gas are examined. These include the properties of nitrogen, particularly at high pressure; isentropic expansion and normal shock flows in nitrogen; real gas ratios; and the problem of condensation. Sources of information on cryogenic technology are cited.

Kilgore, R. A.↗

NASA NTRS1976

Design features and operational characteristics of the Langley 0.3-meter transonic cryogenic tunnel

Experience with the Langley 0.3 meter transonic cryogenic tunnel, which is fan driven, indicated that such a tunnel presents no unusual design difficulties and is simple to operate. Purging, cooldown, and warmup times were acceptable and were predicted with good accuracy. Cooling with liquid nitrogen was practical over a wide range of operating conditions at power levels required for transonic testing, and good temperature distributions were obtained by using a simple liquid nitrogen injection system. To take full advantage of the unique Reynolds number capabilities of the 0.3 meter transonic tunnel, it was designed to accommodate test sections other than the original, octagonal, three dimensional test section. A 20- by 60-cm two dimensional test section was recently installed and is being calibrated. A two dimensional test section with self-streamlining walls and a test section incorporating a magnetic suspension and balance system are being considered.

Kilgore, R. A.↗

NASA NTRS1976

Static force tests of a sharp leading edge delta-wing model at ambient and cryogenic temperatures with a description of the apparatus employed

A sharp leading edge delta-wing model was tested through an angle-of-attack range at Mach numbers of 0.75, 0.80, and 0.85 at both ambient and cryogenic temperatures in the Langley 1/3-meter transonic cryogenic tunnel. Total pressure was varied with total temperature in order to hold test Reynolds number constant at a given Mach number. Agreement between the aerodynamic data obtained at ambient and cryogenic temperatures indicates that flows with leading-edge vortex effects are duplicated properly at cryogenic temperatures. The test results demonstrate that accurate aerodynamic data can be obtained by using conventional force-testing techniques if suitable measures are taken to minimize temperature gradients across the balance and to keep the balance at ambient (warm) temperatures during cryogenic operation of the tunnel.

Kilgore, R. A.↗

NASA NTRS1976

Recent progress on new facilities at the NASA Langley Research Center

A new fan-driven high Reynolds number transonic cryogenic tunnel the National Transonic Facility is being planned for the United States. This tunnel will provide an order of magnitude increase in Reynolds number capability over existing tunnels. Theoretical studies and experience with the Langley 1/3 Meter Transonic Cryogenic Tunnel indicate that the cryogenic concept allows the attainment of full-scale Reynolds number at reasonable levels of dynamic pressure. The unique modes of operation which are available only in a cryogenic tunnel make possible the separation of Mach number, Reynolds number possible the separation of Mach number, Reynolds number, and aeroelastic effects. By reducing the drive power requirements to a level where a conventional fan drive system may be used, the cryogenic concept makes possible a tunnel with high productivity and run times sufficiently long to allow for all types of tests at reduced capital costs and, reduced total energy consumption.

Kilgore, R. A.↗

NASA NTRS1976

The cryogenic transonic wind tunnel for high Reynolds number research

Based on theoretical studies and experience with a low speed cryogenic tunnel and with the transonic cryogenic tunnel, the cryogenic wind tunnel concept has been shown to offer many advantages with respect to the attainment of full scale Reynolds number at reasonable levels of dynamic pressure in a ground based facility. The unique modes of operation available in a pressurized cryogenic tunnel make possible for the first time the separation of Mach number, Reynolds number, and aeroelastic effects.

Kilgore, R. A.↗

NASA NTRS1976

The cryogenic wind tunnel

Based on theoretical studies and experience with a low speed cryogenic tunnel and with a 1/3-meter transonic cryogenic tunnel, the cryogenic wind tunnel concept was shown to offer many advantages with respect to the attainment of full scale Reynolds number at reasonable levels of dynamic pressure in a ground based facility. The unique modes of operation available in a pressurized cryogenic tunnel make possible for the first time the separation of Mach number, Reynolds number, and aeroelastic effects. By reducing the drive-power requirements to a level where a conventional fan drive system may be used, the cryogenic concept makes possible a tunnel with high productivity and run times sufficiently long to allow for all types of tests at reduced capital costs and, for equal amounts of testing, reduced total energy consumption in comparison with other tunnel concepts.

Kilgore, R. A.↗

NASA NTRS1975

Dynamic-stability tests on an aircraft escape module at Mach numbers from 0.40 to 2.16

Wind-tunnel measurements of the aerodynamic damping and oscillatory stability of a model of a proposed escape module for a military aircraft have been made using a small-amplitude forced-oscillation technique in pitch and yaw at Mach numbers from 0.40 to 2.16 and in roll at Mach numbers from 0.40 to 1.20. The results in pitch indicate regions in the angle-of-attack range where the model exhibits large and rapid changes in both damping and stability with angle of attack, probably caused by vortex flow over the fins. There was no pronounced effect of change in angle of attack on damping in yaw. Except for the highest Mach number, negative damping in roll was produced at high negative angles of attack.

Davenport, E. E.↗

NASA NTRS1975

Analysis of validation tests of the Langley pilot transonic cryogenic tunnel

A pilot transonic cryogenic pressure tunnel has recently been developed and proof tested at the NASA Langley Research Center. In addition to providing an attractive method for obtaining high Reynolds number results at moderate aerodynamic loadings and tunnel power, this unique tunnel allows the independent determination of the effects of Reynolds number, Mach number, and dynamic pressure (aeroelasticity) on the aerodynamic characteristics of the model under test. The proof of concept experimental and theoretical studies are briefly reviewed. Experimental results obtained on both two- and three-dimensional models have substantiated that cryogenic test conditions can be set accurately and that cryogenic gaseous nitrogen is a valid test medium.

Ray, E. J.↗

NASA NTRS1975

Aerodynamic damping and oscillatory stability in pitch of a model of a proposed manned lifting entry vehicle at Mach Numbers of 1.80, 2.16, and 2.86

Wind tunnel tests were conducted using a model of a proposed manned lifting entry vehicle to determine the aerodynamic damping and oscillatory stability in pitch. The model was tested at Mach numbers of 1.80, 2.16, and 2.86. Angles of attack varied from minus 2 degrees to plus 30 degrees at zero angle of sideslip using a small-amplitude, forced-oscillation technique. It was determined that, in general, all the configurations have near zero or slightly positive damping in pitch throughout the angle of attack range. The effects of the deflection of flaps on aerodynamic damping are discussed.

Kilgore, R. A.↗

NASA NTRS1975

Aerodynamic damping and oscillatory stability in pitch and yaw of a model of a proposed manned lifting entry vehicle at Mach numbers from 0.20 to 1.20

Wind tunnel tests have been made at angles of attack from about -2 deg to about 22 deg at 0 deg angle of sideslip by using a small-amplitude forced-oscillation technique. Models were tested with upper and lower control flaps both deflected and undeflected. The configuration with flaps deflected has positive damping in both pitch and yaw and is stable in both pitch and yaw except at the higher angles of attack where the tail surfaces are submerged in the wake from the body.

Kilgore, R. A.↗

NASA NTRS1975

Cryogenic nitrogen as a transonic wind-tunnel test gas

The test gas for the Langley Pilot Transonic Cryogenic Tunnel is nitrogen. Results from analytical and experimental studies that have verified cryogenic nitrogen as an acceptable test gas are reviewed. Real-gas isentropic and normal-shock flow solutions for nitrogen are compared to the ideal diatomic gas solutions. Experimental data demonstrate that for temperatures above the liquefaction boundaries there are no significant real-gas effects on two-dimensional airfoil pressure distributions. Results of studies to determine the minimum operating temperatures while avoiding appreciable effects due to liquefaction are included.

Adcock, J. B.↗

NASA NTRS1974

Simulation of flight test conditions in the Langley pilot transonic cryogenic tunnel

The theory and advantages of the cryogenic tunnel concept are briefly reviewed. The unique ability to vary temperature independently of pressure and Mach number allows, in addition to large reductions in model loads and tunnel power, the independent determination of Reynolds number, Mach number, and aeroelastic effects on the aerodynamic characteristics of the model. Various combinations of Reynolds number and dynamic pressure are established to represent accurately flight variations of aeroelastic deformation with altitude changes. The consequences of the thermal and caloric imperfections of the test gas under cryogenic conditions were examined and found to be insignificant for operating pressures up to 5 atm. The characteristics of the Langley pilot transonic cryogenic tunnel are described and the results of initial tunnel operation are presented. Tests of a two-dimensional airfoil at a Mach number of 0.85 show identical pressure distributions for a chord Reynolds number of 8,600,000 obtained first at a stagnation pressure of 4.91 atm at a stagnation temperature of 322.0 K and then at a stagnation pressure of 1.19 atm at a stagnation temperature of 116.5 K.

Kilgore, R. A.↗