Engineering topics
Publications and source records attributed to Goodyer, M. J..
The techniques used to find aerodynamically straight wall contours in a test section of a transonic wind tunnel are discussed. The walls were defined as aerodynamically straight up to Mach 0.9.
An improvement is presented for the 2-D strategies for adjustment of the flexible top and bottom walls of an Adaptive (Wind Tunnel) Wall Test Section (AWTS). This adjustment is part of the wall adaptation process to eliminate top and bottom wall interference at the source. The improvements to account for second order effects are described in mathematical detail. It is intended that these improvements should further minimize the necessary iterations in the wall adaptation process. An associated computer program written in BASIC is presented and several test cases run with this program are discussed. The strategy performs well for a theoretical test case but when applied to experimental AWTS data some discrepancies in the adapted wall shapes are found.
Data for a NACA 0012-64 airfoil at Mach numbers of 0.3, 0.5, and 0.7 with the test section flexible walls set in the "straight" position are presented because they may prove useful in the development of wind tunnel correction techniques.
This design eliminates the uncertainties in data from conventional transonic test sections. Sidewalls are rigid, and the flexible floor and ceiling are positioned by motorized jacks controlled by on-line computer to minimize run times. The tunnel-computer combination is self-streamlining without reference to the model. Data is taken from the model only when the walls are good streamlines, and is corrected for the small, known but inevitable residual interferences. Two-dimensional validation testing in the Mach range up to about 0.85 where the walls are just supercritical shows good agreement with reference data using a height:chord ratio of 1.5. Techniques are under development to extend Mach number above 1. This work has demonstrated the feasibility of almost eliminating wall interferences, improving flow quality, and reducing power requirements or increasing Reynolds number. Extensions to three-dimensional testing are outlined.
The technology which is required to allow the principles of magnetic suspension and balance systems (MSBS) to be applied to the high Reynolds number transonic testing of aircraft models is examined. A test facility is presented as comprising a pressurized transonic cryogenic wind tunnel, with the MSBS providing full six degree of freedom control. The electro-magnets which are superconducting and fed from quiet, bipolar power supplies are examined. A model control system having some self adaptive characteristics is discussed.
The characteristics and behavior of a 0.3m transonic cryogenic wind tunnel are discussed. The wide band of usable Reynolds numbers is analyzed along with a flow visualization technique using propane. The combination of magnetic suspension with the cryogenic wind tunnel is described. An outline of the circuit showing the locations of the magnet system and the features of the tunnel are presented.
The evolution in Europe of the flexible walled test section, as applied to two dimensional testing at low and transonic speeds, is traced from its beginnings at NPL, London, in the early 1940's, and is shown to lead logically to the latest version now nearing completion at Southampton University. The principal changes that have taken place are improvements in the methods of choosing wall contours such that they rapidly follow appropriate streamlines, and reductions in the depth of test sections. The latest transonic test section presently under assembly at Southampton has, as its principal new feature, the facility for the automation of wall streamlining with the aid of an on-line computer. The versatility of the flexible walled test section is emphasized by reference to the simulation of alternative flows including cascade, steady pitching in an infinite flowfield, and ground effect. Finally, sources of error in streamlining are identified, with methods for their alleviation.
The design and performance of a low speed one atmosphere cryogenic wind tunnel is described. The tunnel is fan driven and operates over the temperature range 305K to 77K at Mach numbers up to 0.28. It is cooled by the injection and evaporation of liquid nitrogen in the circuit, and the usual test gas is nitrogen. The tunnel has a square test section 0.1m across and was built to allow, at low costs, the development of testing techniques and the development of instrumentation for use in cryogenic tunnels, and to exploit in general instrumentation work the unusuallly wide range of unit Reynolds number available in such tunnels. The tunnel was first used in the development of surface flow visualization techniques for use at cryogenic temperatures.
Testing in the low speed flexible walled tunnel in an effort to explain the reasons for data discrepancies at high angles of attack are presented. Automated transonic test sections were developed. The flexible walled tunnel was used in a new operating mode to a generated curved flow around the airfoil, allowing the extraction of purely rotary derivatives. Some straight wall, low speed pressure data, for wall and model, which is used for checking interference correction methods were reported. Computer software which includes an old streamlining algorithm and a prediction algorithm was examined.
Comprehensive aerodynamic data on an airfoil section were obtained through a wide range of angles of attack, both stalled and unstalled. Data were gathered using a self streamlining wind tunnel and were compared to results obtained on the same section in a conventional wind tunnel. The reduction of wall interference through streamline was demonstrated.
It is shown that certain two dimensional bodies exist having a sensitivity of surface pressure to flow direction which is much higher than for currently used two dimensional yawmeters. A body having a simple cross sectional shape was selected for experimental evaluation at low Mach number, but at a typical Reynolds number for such instruments.
A method has been developed for accurately computing the imaginary flow fields outside a flexible walled test section, applicable to lifting and non-lifting models. The tolerances in the setting of the flexible walls introduce only small levels of aerodynamic interference at the model. While it is not possible to apply corrections for the interference effects, they may be reduced by improving the setting accuracy of the portions of wall immediately above and below the model. Interference effects of the truncation of the length of the streamlined portion of a test section are brought to an acceptably small level by the use of a suitably long test section with the model placed centrally.
A two dimensional test section in a low speed wind tunnel is producing flow conditions free from wall interference. The test section has flexible top and bottom walls, and rigid sidewalls from which the models are mounted spanning the tunnel. All walls are unperforated, and the flexible walls are positioned by screw jacks. To eliminate wall interference, the wind tunnel itself supplies the information required in this streamlining process, when run with the model present. Measurements taken at the flexible walls are used by the tunnel computer to check wall contours. When the static pressure distribution in the test section along a contoured flexible wall matches that computed for an imaginary flow field passing over the outside of the same contour, the wall is a streamline in an infinite flow field and the test section flow is free from wall interference. A series of iterations brings the walls from straight to streamlines. Illustrative aerodynamic data is presented, taken on a bluff body and a lifting wing.
A two dimensional test section in a low speed wind tunnel capable of producing flow conditions free from wall interference is presented. Flexible top and bottom walls, and rigid sidewalls from which models were mounted spanning the tunnel are shown. All walls were unperforated, and the flexible walls were positioned by screw jacks. To eliminate wall interference, the wind tunnel itself supplied the information required in the streamlining process, when run with the model present. Measurements taken at the flexible walls were used by the tunnels computer check wall contours. Suitable adjustments based on streamlining criteria were then suggested by the computer. The streamlining criterion adopted when generating infinite flowfield conditions was a matching of static pressures in the test section at a wall with pressures computed for an imaginary inviscid flowfield passing over the outside of the same wall. Aerodynamic data taken on a cylindrical model operating under high blockage conditions are presented to illustrate the operation of the tunnel in its various modes.
Theoretical considerations indicate that cooling the wind-tunnel test gas to cryogenic temperatures will provide a large increase in Reynolds number with no increase in dynamic pressure while reducing the tunnel drive-power requirements. Studies were made to determine the expected variations of Reynolds number and other parameters over wide ranges of Mach number, pressure, and temperature, with due regard to avoiding liquefaction. Practical operational procedures were developed in a low-speed cryogenic tunnel. Aerodynamic experiments in the facility demonstrated the theoretically predicted variations in Reynolds number and drive power. The continuous-flow-fan-driven tunnel is shown to be particularly well suited to take full advantage of operating at cryogenic temperatures.
A method and apparatus for measuring the stagnation pressure of supersonic velocity gas streams without the generation of shock waves which interfere with such measurements are given. The technique is insensitive to the type of gas and Mach number and is therefore particularly useful in the study of jet engine exhausts.
Probe was developed which is capable of recording true stagnation pressure in supersonic and subsonic flows. Static pressure orifices in the compression surface allow probe to be used as combined flow direction and stagnation pressure sensor.
Theoretical considerations indicate that cooling the wind-tunnel test gas to cryogenic temperatures will provide a large increase in test Reynolds number with no increase in dynamic pressure while reducing the tunnel drive-power requirements. Studies have been made to determine the expected variations of Reynolds number and other parameters over wide ranges of Mach number, pressure, and temperature with due regard to avoiding liquefaction and adverse real-gas effects. Practical operational procedures have been developed in a low-speed prototype cryogenic wind tunnel. Aerodynamic experiments in the facility have demonstrated the theoretically predicted variations in Reynolds number and drive power. Force and moment measurements on a wing model mounted on a water-jacketed strain-gage sting balance have demonstrated the feasibility of operation of such balances in a cryogenic environment.