The High Temperature Gas Dynamics Facility was developed as a result of the Aeronautical Systems Division's effort to extend the state-of the-art in hypersonic aerodynamic simulation. The facility is a hypersonic wind tunnel supplied with high pressure air, heated from a zirconia dioxide pebble heater. The maximum stagnation pressure and temperature is 40 atres and 4700 R, respectively. This facility is one of four of its kind in this hemisphere and the only Air Force facility of its type. This report discusses the modification of the facility to a two foot diameter test section with a Mach number range of 6 through 14 and its expected performance. This facility is scheduled to be operational in the Fall of 1963.

The High Temperature, Hypersonic Gasdynamics Facility was developed as a result of the efforts of the Aeronautical Systems Division to extend the state-of-the art in hypersonic aerodynamics simulation. The High Temperature Facility is an operational hypersonic wind tunnel supplied by high pressure, heated air from a zirconium oxide pebble heater. The maximum stagnation pressure is 40 atmospheres, and the maximum temperature is 4700 R. The facility is one of four of its kind in this hemisphere and the only Air Force facility of its type. This report describes its operation from August 1961 until May 1962 as a Mach 4 aerodynamic test facility. The successful operation of this facility is a significant achievement in the area of hypersonic aero dynamic testing techniques.

An experimental investigation of the test section flow blockage characteristics was conducted at Mach 4 in the High Temperature Hypersonic Gas dynamics Facility. The models utilized in this program were pointed and blunted cones from 5 degree half angle to 90-degree half angle, flat plates, delta winged shapes, and hemispherical models, some of which were run up to 40-degree angle of attack. Comparison of the data with other facilities resulted in a correlation of the maximum model size compared to the potential flow core size with Mach number and drag coeffi cient for both open and closed test section con figurations for Mach numbers from 1.5 to greater than 10.

The purpose of this test was to develop diagnostic techniques for the detection of hydrogen combustion in a hypersonic high temperature air stream. This was accomplished by ejecting hydrogen through sonic nozzles oriented normal to the airstream in the case of the flat plate model and at a 45-degree angle downstream in the case of the stepped flat plate model. The hydrogen gas was ejected into a high temperature airstream (T sub 0 4000 R) at Mach Number 4.14 to 4.24. The data is presented in three forms: (1) photographs; (2) graphical; and (3) oscillo graph traces.

Assuming the wall to have zero heat capacity and conductivity, an energy balance is written across the surface of a vehicle equating the aerodynamic heat transferred to the wall to the heat radiated from the wall surface. Laminar and turbulent flow are considered over the following shapes: axisymmetric, 2-dimensional leading edge, flat plate, and pointed cone. The energy balance equation is reduced in terms of Tw and parameters for the above surface shapes. Graphical solutions of the resulting equations are presented allowing Tw to be found at any point on the above surfaces. An additional relationship is presented, along with its associated plots, allowing the heating rate to the surface to be found at any point.