Abstract:
The invention disclosed herein provides a subsonic wind tunnel capable for accurately maintaining a desired Mach number, pressure and temperature for use in aerothermal testing of materials. The invention allows the Mach number to be precisely controlled in the test section of the wind tunnel by employing a restricted outlet acting as a sonic throat for the wind tunnel. In the preferred embodiment, the restricted outlet is constructed to be variable in cross sectional area allowing a range of Mach numbers to be tested. The variable outlet is varied during the operation of the wind tunnel so that an actual trajectory with changing Mach numbers, pressure, and temperature is simulated.

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
   The invention described herein may be manufactured and used by or for the government of the United States of America for governmental purposes without the payment of any royalties thereon or therefore. 
   BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   The present invention relates to a wind tunnel testing apparatus and methods for use in aerothermal testing, and more particularly to such materials and methods utilizing a low velocity subsonic flow with shear and heat flux conditions equivalent to a hypersonic or supersonic flight path of a test object. 
   2. Description of the Prior Art 
   The use of wind tunnels for simulating flight conditions is a common practice in the aerospace industry. Utilizing wind tunnels, objects are subjected to carefully controlled environmental conditions and monitored to determine how the object behaves. All wind tunnels operate under the same basic principle. A gas is accelerated to a desired velocity with a test object located within the gas stream. The conditions in the wind tunnel are controlled for a variety of parameters such as pressure, temperature, and velocity. 
   Wind tunnels are typically categorized depending on the speed of the gas flow that they produce. The gas speed is characterized by a Mach number, which is the speed of the gas divided by the speed of sound in the gas. Wind tunnels are generally divided into one of four categories depending on the Mach number of the gas flow: subsonic, transonic, supersonic, and hypersonic. Subsonic wind tunnels are the simplest of the four and are relatively inexpensive. They accelerate a gas to a speed less than that of sound, usually up to Mach 0.6. Increasing in complexity, but still relatively simple are transonic wind tunnels that provide for gas flows near the speed of sound, from Mach 0.6 up to Mach 0.9. Supersonic wind tunnels accelerate the gas beyond the speed of sound, over Mach 1, and hypersonic wind tunnels accelerate the gas to well beyond Mach 2. As the required gas velocity increases, the wind tunnel becomes increasingly more expensive to build and operate. 
   Traditionally, an object to be tested is subjected to environmental conditions that are similar to the environmental conditions the object will experience in normal operation. In most applications, this requires that the gas in the wind tunnel be heated and accelerated to a significant fraction of the flight speed of the test object. Additionally, to simulate high altitude conditions, the ambient pressure in the wind tunnel may need to be reduced. 
   Using present technology to test materials for a missile designed to fly at Mach 6 would require the test gas to be heated to 2500 F and the gas to be accelerated to 6 times the speed of sound. This has several drawbacks. First, in order to accelerate the gas to such a high velocity, a converging-diverging nozzle must be used. A converging-diverging nozzle is a nozzle that accelerates the flow of subsonic gas to the speed of sound as it converges at the throat, or minimum diameter of the nozzle, and then, as the nozzle diverges, the gas is further accelerated to supersonic and hypersonic speeds. Such nozzles are expensive to build and a separate nozzle is required for each Mach number of interest. The nozzle must be as large as, if not larger, than the object to be tested. In addition to being expensive to build, massive amounts of gas must be available to supply the high mass flow rates encountered with such a large nozzle. Finally, the nozzles themselves must be kept cool, thus requiring large cooling systems. These limitations restrict the practical size of the nozzle commonly resulting in use of a nozzle that is undersized for a given application. 
   Aerothermal testing examines how a material responds to conditions of high temperatures and viscous shear forces. In applications that are not shear sensitive, matching the heat flux to that expected under operational flight conditions is sufficient. Testing other materials, such as ablators, requires that both the heat flux and shear force must be simultaneously matched. The parameters that primarily contribute to the heat flux and shear force are the pressure, temperature, and Mach number of the gas to which the test object is subjected. For any given point on a high-speed flight trajectory, an equivalent subsonic flow condition having a different pressure and temperature exists that will provide matching heat flux and shear conditions. The required values of pressure, temperature, and Mach number are calculated using conventional methods such as standard closed-form empirical boundary layer approximations, computational fluid dynamics, or other computational programs such as ATAC available from ITT Aerotherm. 
   Prior attempts at simulating high-speed aerothermal conditions with a subsonic test section relied on a diffuser from which the test gas exited directly into atmospheric pressure or into a controlled pressure environment. With this arrangement, it was not possible to control the test section pressure or Mach number and hence it was only possible to simulate the heat flux over the test article. Other problems were encountered due to a mismatch between the required mass flow and the scale of the diffuser and resulted in a non-uniform flow condition over the test article. 
   SUMMARY OF THE INVENTION 
   The present invention solves the problem of having to accelerate a hot gas to supersonic or hypersonic speeds to accurately perform aerothermal testing upon a test object under conditions simulating supersonic or hypersonic flight. The present invention permits a subsonic condition to be utilized to match the shear and heat flux that would occur during a supersonic or hypersonic flight trajectory. The Mach number, pressure, and temperature in the wind tunnel are all easily controllable to match a predetermined heat flux and shear. 
   The invention permits larger items to be tested in a given wind tunnel because lower mass flow rates are required. Utilizing the present invention, the mass flow rates required to test the object are greatly reduced as compared to a hypersonic wind tunnel. For an equivalent mass flow rate, the present invention affords the use of a larger test section allowing for testing of larger objects than would otherwise be the case. 
   A feature of the present invention is that the static pressure surrounding the test article need not be reduced to simulate high altitude, low-pressure flight. Rather, the operating pressure may be chosen initially and then equivalent subsonic conditions determined to match that pressure. It is beneficial to be able to choose the operating pressure because as the operating pressure is increased, the Mach number required for the equivalent subsonic condition decreases. 
   The invention additionally permits the Mach number in the test section to be varied during operation of the wind tunnel. Conventional technology typically is limited to testing at a single condition. An accurate simulation of a complete trajectory requires conditions that change in time to match those experienced in flight. The present invention allows the heat flux and shear to be varied in time to accurately simulate the complete trajectory of a test article. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a general block diagram of a subsonic wind tunnel facility. 
       FIG. 2  is a schematic of a subsonic diffuser and test section fitted with a restricted area outlet throat. 
       FIG. 3  is a detailed schematic of a test item. 
       FIG. 4  is a view of a variable area outlet throat in a substantially open position. 
       FIG. 5  is a view of a variable area outlet throat in a partially restricted position. 
       FIG. 6  is a view of a variable area outlet throat in a substantially restricted position. 
   

   DESCRIPTION OF THE PREFERRED EMBODIMENTS 
   Referring to  FIGS. 1 &amp; 2 , subsonic wind tunnel  10  is described. Air compressor  12  provides compressed air  13  to compressed air storage tank  14 . Compressed air storage tank  14  delivers compressed air  13  to sudden expansion air heater  20  where compressed air  13  is mixed with fuel  15  such as propane supplied by fuel storage tank  16 . It will be understood that other conventional sources of compressed air, such as fans, turbines and the like may also be used to supply compressed air. Hydrogen  17  is optionally added to the fuel air mixture to enhance ignition of the fuel air mixture. Alternatively, compressed air  13  is heated using other conventional methods such as a vitiated air heater, arc jet, or pebble bed heater to produce heated gas  32 . Pressurized water supply  22  cools wind tunnel  10  and helps regulate the temperature of the wind tunnel components. Heated gas  32  flows through flow straightener  31  and into diffuser lumen  25  of subsonic diffuser  26 . Subsonic diffuser  26  is coupled to test section  28  having test section lumen  27  therethrough that is adapted to hold test object  30 . Heated gas  32  flows through test section lumen  27 , around test object  30  and out the outlet throat  40  of test section  28 . 
   Air compressor  12  delivers compressed air  13  at high pressure to be stored within compressed air storage  14 . Compressed air storage  14  provides a continuous supply of compressed air  13  at a mass flow rate that is capable of exceeding the output capacity of air compressor  12 . The total amount of compressed air  13  available in compressed air storage  14  is typically the limiting factor in the duration of operation of wind tunnel  10 . Compressed air storage  14  is connected or coupled to sudden expansion air heater  20  through air control valve  19  that allows the mass flow rate of compressed air  13  to be controlled. 
   Fuel storage  16  is connected or coupled to sudden expansion air heater  20  through fuel control valve  21  which regulates the mass flow rate of fuel  15  delivered to sudden air expansion heater  20 . Combustion of fuel  15  and compressed air  13  is optionally facilitated by the introduction of hydrogen  17  provided by hydrogen storage  18 . Once the fuel air mixture is ignited, hydrogen  17  is turned off at hydrogen control valve  23 . The combustion of compressed air  13  and fuel  15  produces heated gas  32 . The mass flow rate of fuel  15  determines the amount of heat added to wind tunnel  10  and is easily adjusted using fuel control valve  21 . 
   Water lines  24  conduct water from pressurized water supply  22  to sudden expansion air heater  20  to provide for cooling. The water helps to keep wind tunnel  10  from overheating and is additionally used to control the operating temperature of the wind tunnel components. 
   In operation, the mixture of the three gases is ignited within sudden expansion air heater  20 . The total mass flow rate of the resulting heated gas  32  is controlled using control valves  19  and  21  from compressed air storage  14  and fuel storage  16 , respectively. The temperature of heated gas  32  is regulated by adjusting the mass flow rate of fuel  15  entering air heater  20 . The mass flow rate of fuel  15  is controlled utilizing feedback from a gas temperature sensor preferably disposed in test section lumen  27 . Heated gas  32  is thoroughly mixed within sudden expansion gas heater  20  before exiting. Heated gas  32  exiting sudden expansion air heater  20  flows into flow straightener  31  providing a more uniform and less turbulent flow. 
   Heated gas  32  flows through flow straightener  31  and into subsonic diffuser lumen  25 . Proceeding along the direction of gas flow, the cross sectional area of subsonic diffuser lumen  25  gradually increases from the initial cross section at its coupling or juncture with flow straightener  31  until the cross section of subsonic diffuser lumen  25  matches the cross section of test section lumen  27  at the coupling or juncture of subsonic diffuser lumen  25  and test section lumen  27 . As heated gas  32  flows through subsonic diffuser lumen  25  its Mach number decreases and its pressure increases. 
   The larger end of diffuser lumen  25  opposite flow straightener  31  opens into and is coupled to test section lumen  27  located within test section  28  of wind tunnel  10 . Test section lumen  27  is generally cylindrical in shape with the opening from diffuser lumen  25  on one end and an outlet opening located at the opposite end. Test section lumen  27  is adapted to removably retain test object  30 . Test object  30  has an object test region  36  that is generally cylindrical in shape. The center longitudinal axis YY of test object  30  is aligned with center longitudinal axis ZZ of test section lumen  27  so that the distance from the surface of cylindrical test object  30  to the walls of test section lumen  27  preferably is constant circumferentially. The cross sectional area of the resulting test section lumen  27  space located between the inner wall of test section  28  and the outer surface of test object  30  is constant over the longitudinal extent of object test section  36 . 
     FIG. 3  depicts a representative test object  30  detailing three regions of test object  30 . The region of test object  30  first encountered by hot gas flow  32  is nose region  34 . Nose region  34  is water cooled to prevent overheating and to minimize influencing the measurements taken within object test region  36 . Object test region  36  is the area where materials are actually tested. In a typical representative example, a material to be tested is applied or disposed circumferentially in this region of test object  30 . Behind or downstream of test region  36  is aft region  38 , which, like nose region  34 , is cooled to prevent overheating and to minimize influencing the measurements taken within object test region  36 . 
   Disposed in and coupled to test section lumen  27  is a variable area outlet throat  40  having a cross sectional area that is smaller than the cross sectional area of any gas fillable region (i.e., that region not occupied by test articles and related apparatus) within wind tunnel  10  and downstream of sudden expansion air heater  20 . The variable area outlet throat  40  provides a restricted opening available for gases to exit from wind tunnel  10 . 
     FIGS. 4 ,  5 , and  6  depict a representative example of a variable area outlet throat  40 . The variable area outlet throat  40  is formed by proximal plate  42  and a distal plate  44 . Plates  42  and  44  are flush with each other and axially aligned with the test section lumen  27 . Proximal plate  42  is coupled to the test section  28 . Distal plate  44  is coupled to the test section  28  but allowed to rotate freely around the longitudinal axis YY test object  30 . Plates  42  and  44  have circumferentially aligned passages allowing fluid to flow therethrough.  FIG. 4  depicts outlet throat in a substantially open position. The passages of plates  42  and  44  are aligned, and outlet throat  40  area is equal to the size of the passages.  FIG. 5  depicts outlet throat  40  in a partially restricted position. In the partially restricted position, distal plate  44  has been rotated so that the passages are no longer aligned. The resulting outlet throat  40  is now the passage formed by the overlapping area of passages through plates  42  and  44 .  FIG. 6  depicts outlet throat in a substantially restricted position. In the substantially restricted position, distal plate  44  has been further rotated until there is almost no overlap between plates  42  and  44 . The outlet throat  40  area is calibrated to provide a determined cross sectional area for a given rotation of the distal plate  44 . 
   At low mass flow rates of heated gas  32 , the flow will be subsonic through outlet throat  40 . As heated gas  32  mass flow rate is increased, heated gas  32  Mach number in outlet throat  40  will increase until the Mach number of heated gas  32  reaches one. Further increases of heated gas  32  mass flow rate have no effect on heated gas  32  Mach number in outlet throat  40  because the flow is choked by a sonic throat formed by heated gas  32  in outlet throat  40 . 
   So long as the heated gas  32  mass flow rate is at or above the rate resulting in choked flow of heated gas  32  within outlet throat  40 , the cross sectional area of outlet throat  40  controls the Mach number of heated gas  32  flowing through test section lumen  27 . The Mach number of heated gas  32  flowing through test section lumen  27  is a function of the ratio of the cross sectional areas of outlet throat  40  and of the gas fillable region of test section lumen  27 . Reference to an accepted compressible subsonic flow table available in most fluid dynamics references will supply the corresponding Mach numbers for each cross sectional area ratio. The Mach number of heated gas  32  flowing in test section lumen  27  and through outlet throat  40  is independent of gas pressure or temperature. Rather, the Mach number of heated gas  32  depends solely upon the ratio of the cross sectional areas of test section lumen  27  and outlet throat  40 . 
   Wind tunnel  10  is designed to test the aerothermal properties of a material applied to test region  36  of test object  30 . The aerothermal properties of interest are the shear and heat flux experienced during flight conditions. These properties are primarily dependent upon the Mach number, the pressure and the temperature of the gas flowing past the material. The heat flux and shear experienced during flight is measured and recorded or the values are calculated using common computational techniques. 
   For each supersonic or hypersonic flight condition to be tested, a corresponding subsonic condition exists that will result in an equivalent shear and heat flux. 
   As has been described, the present invention provides for simple selection and regulation of the temperature, Mach number, and pressure of heated gas  32  flowing through test section lumen  27 . The Mach number of heated gas  32  flowing in test section lumen  27  is controlled by varying the cross sectional area of outlet throat  40 . Alternatively, the Mach number within test section lumen  27  is monitored, while the size of the cross section of outlet throat  40  is varied, until the desired heated gas  32  Mach number is obtained. Using either method, once the heated gas  32  Mach number to be used for a test is selected by adjusting the cross sectional area of outlet throat  40 , the Mach number of heated gas  32  flowing through test section lumen  28  will remain constant regardless of fluctuations in gas temperature, gas pressure, or gas mass flow. The temperature of heated gas  32  in wind tunnel  10  is adjusted by changing the amount of heat entering the wind tunnel  10 . This is controlled by adjusting the amount of fuel  15  that is introduced into sudden expansion air heater  20 . It is possible to precalculate the rate of fuel consumption required to give a certain chosen temperature, or the rate is adjusted empirically until the desired temperature is realized. 
   The gas pressure in test section lumen  27  is controlled by adjusting the mass flow rate of heated gas  32  entering wind tunnel. Each of the three gases (compressed air  13 , fuel  15 , and hydrogen  17 ) is individually controlled for mass flow rate and the total mass flow rate is the combined mass flow rates of all three. Because hydrogen  17  contributes a negligible amount to the overall mass flow rate, only fuel  15  and compressed air  13  need be considered. The standard equation of state readily provides a total mass flow that will result in the desired pressure in test section lumen  27 . 
   In operation, wind tunnel  10  provides heated gas  32  flow at a controlled Mach number, pressure, and temperature to test a material under conditions simulating those of a trajectory flown by a vehicle that may incorporate the material. This is accomplished by varying the cross sectional area of variable area outlet throat  40  in real time during a test. As the area of outlet throat  40  is varied, the Mach number and pressure of heated gas  32  in test section lumen  27  will change. The mass flow rate of heated gas  32  entering system  10  is adjusted independently of, or together with variable area outlet throat  40  to maintain or to vary the pressure in a desired profile. The temperature of heated gas  32  is varied to match a desired profile at the same time. 
   While the present invention has been described in connection with what are currently considered to be the most practical and preferred embodiments, it is to be understood that the invention is not to be limited to the disclosed embodiments, but to the contrary, is intended to cover various modifications, embodiments, and equivalent processes included within the spirit of the invention as may be suggested by the teaching herein, which are set forth in the appended claims, and which scope is to be accorded the broadest interpretation so as to encompass all such modification, embodiments, and equivalent processes.