Abstract:
The present invention blocks and/or attenuates the upstream travel of acoustic disturbances or sound waves from a flight vehicle or components of a flight vehicle traveling at subsonic speed using a local injection of a high molecular weight gas. Additional benefit may also be obtained by lowering the temperature of the gas. Preferably, the invention has a means of distributing the high molecular weight gas from the nose, wing, component, or other structure of the flight vehicle into the upstream or surrounding air flow. Two techniques for distribution are direct gas injection and sublimation of the high molecular weight solid material from the vehicle surface. The high molecular weight and low temperature of the gas significantly decreases the local speed of sound such that a localized region of supersonic flow and possibly shock waves are formed, preventing the upstream travel of sound waves from the flight vehicle.

Description:
STATEMENT OF GOVERNMENT INTEREST 
     The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without payment of any royalties thereon or therefor. 
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates generally modifying the local speed of sound characteristics of a gas or fluid flow, more particularly to modifying these characteristics using the distribution of a high molecular weight and/or low temperature gas into to the flow, and most particularly to attenuating the acoustic disturbance and/or noise associated with aircraft by modifying the local speed of sound characteristics using high molecular weight and/or low temperature gas upstream of the aircraft. 
     2. Description of the Related Art 
     Reduction of noise generated by jet engines has become an important aspect of aircraft design. Sound waves from aircraft traveling at subsonic speed propagate in all directions relative to the aircraft. The reduction of noise generated by aircraft has obvious benefits, including reducing noise pollution and, thereby, reducing safety risks to persons in the vicinity of aircraft. Also, sound waves have a tendency to contribute to wear and fatigue of aircraft components. 
     Many technical solutions have been employed in order to reduce noise generated by aircraft. These include using physical barriers attached to aircraft in order to provide a physical shield to block or attenuate sound waves and injecting gases directly into the jet engine exhaust in order to suppress noise. These solutions provide some reduction to noise levels, but normally result in some reduction in aircraft efficiency or performance. In addition, all of these solutions relate to decreasing sound waves only “downstream” from the aircraft, but do not address sound waves emanating “upstream” from the aircraft. 
     Therefore, it is desired to provide a relatively low-cost apparatus and method to attenuate or block aircraft noise emanating in any direction from an aircraft, without affecting aircraft performance. 
     SUMMARY OF THE INVENTION 
     The invention proposed herein comprises a device and method for blocking an/or attenuating the upstream movement of acoustic disturbances or sound waves emanating from an aircraft flying at subsonic speeds. By blocking the acoustic disturbances, the invention eliminates or reduces the sound perceived by an approaching aircraft. 
     Accordingly, it is an object of this invention to provide a method to block or attenuate acoustic disturbances or sound waves emanating from an aircraft flying at subsonic speeds. 
     It is a further object of this invention to provide a shield for acoustic disturbances or sound waves emanating from an aircraft that does not significantly degrade aircraft performance. 
     It is yet a further object of this invention to provide a method to reduce noise from an aircraft that is relatively low-cost. 
     This invention meets these and other objectives related to reducing aircraft sound by providing a device and method to block the upstream travel of sound waves by the formation of local supersonic flow and possibly shock waves due to an increase in the local speed of sound. By distributing a high molecular weight, and preferably, low temperature gas into the air flow upstream or around an aircraft or flight vehicle, the local speed of sound is greatly decreased, thereby producing local areas of supersonic flow. Shock waves form upstream and around the aircraft structure in these pockets of supersonic flow. The acoustic disturbances and/or sound waves from the flight vehicle can be blocked and/or attenuated with the appropriate positioning of gas injection locations. 
     In addition, the pockets of supersonic flow and local shock waves alter the pressure distribution around the aircraft. This change in the pressure distribution leads to changes in the aerodynamic and stability and control characteristics of the aircraft. This could be used advantageously to impart forces and moments to the aircraft, providing an alternative method of controlling the aircraft without the use of conventional, moving control surfaces. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In the drawings, which are not necessarily to scale, like or corresponding parts are denoted by like or corresponding reference numerals. 
         FIG. 1  depicts an aircraft employing the present invention. 
         FIG. 2   a  depicts a portion of an aircraft employing a direct injection gas system embodiment of the present invention. 
         FIG. 2   b  depicts a portion of an aircraft employing a transverse injection gas system embodiment of the present invention. 
         FIG. 2   c  depicts a portion of an aircraft employing a sublimation system embodiment of the present invention. 
         FIG. 3  depicts changes to the local airflow around a portion of an aircraft employing the present invention. 
         FIG. 4  depicts the movement of acoustic waves from a portion of an aircraft when the invention is both employed and also not employed. 
         FIG. 5  depicts a representation of a computational fluid dynamic analysis model of the invention. 
         FIG. 6  depicts the mass fraction distribution of gas transversely injected into the subsonic flow around an aircraft. 
         FIG. 7  depicts the Mach number distribution with direct gas injection around an aircraft. 
     
    
    
     DESCRIPTION OF PREFERRED EMBODIMENTS 
     The invention, as embodied herein, comprises a device and method for blocking and/or attenuating acoustic waves emanating from aircraft. The invention also provides a method for controlling an aircraft by altering the pressure distribution around the aircraft. 
     In general, the invention provides for distributing a high molecular weight gas (“HMW gas”) into the airflow adjacent/around the aircraft. As used herein, the term “high molecular weight gas” means a gas having a molecular weight of above 100. The HMW gas creates a local area or pocket wherein the speed of sound is reduced, compared to the surrounding airflow. Due to the reduction of the speed of sound in the local area, local areas of supersonic flow, and possibly shock waves, will be formed. Acoustic disturbances or sound waves emanating from the aircraft will be blocked or attenuated by the local areas of supersonic flow. The HMW gas may be distributed into the airflow at a position, or multiple positions, where the user determines it is appropriate to block acoustic waves emanating from the aircraft. 
     In addition to blocking acoustic waves, the local areas of supersonic flow and local shock waves alter the pressure distribution around the aircraft. By altering the pressure distribution around aircraft, changes in the aerodynamic and stability and control characteristics of the aircraft occur. Thus, by changing these characteristics, forces may be imparted to portions of the aircraft in order to alter the aircraft&#39;s direction. 
     Referring to  FIGS. 1-4 , the invention comprises one or more gas distributors  100  positioned on a structure  102  of an aircraft  104  traveling at subsonic speeds. The gas distributors  100  distribute a HMW gas  106  to an area where the user desires to provide reduction in noise created by the aircraft  104 . The HMW gas  106  decreases the speed of sound in a local area  108  where it is distributed. The size of the local area  108  is dependant upon the amount and speed at which the HMW gas  106  is distributed, which may be selected by one skilled in the art consistent with the disclosure herein. 
     In operation, the HMW gas  106  is distributed to form a local area  108  of supersonic flow and possibly shock waves due to a decrease in the local area speed of sound. Acoustic waves  110  emanating from the aircraft  104  are blocked or attenuated by the local area  108  of supersonic flow and shock waves. 
     The gas distributors  100  may employ any mechanism that provides for distributing the HMW gas  106  to create local areas  108  of supersonic flow sufficient in size to provide the desired attenuation of noise from portions of or from an entire aircraft  104 . Two preferred gas distributor  100  mechanisms are an injection system  212  (shown in  FIGS. 2   a  and  2   b ) and a sublimation system  214  (shown in  FIG. 2   c ). 
     One preferred injection system  212  is a direct injection system as shown in  FIG. 2   a . For the direct gas injection system, typical components include a gas storage tank  215 , a gas injection valve  216 , and a gas injector  218 . In this system, the HMW gas  106  is stored at high pressure in the gas storage tank  214  and the gas injection valve  216  is opened to feed the HMW gas  106  through the gas injector  218  into the air flow. Gas injection pressures may be controlled by a regulator, thus controlling the gas penetration distance and dispersion. 
     Another preferred injection system  212  is a transverse injection system as shown in  FIG. 2   b . The components of a transverse injection system include the same components described above for a direct injection system except the gas injector  218  is replaced by an extended boom  220  that extends from the structure  102  and distributes the HMW gas  106  through one or more injectors  222  at the end of the boom  220 . 
     A preferred sublimation system  214  is shown in  FIG. 2   c . This system includes a coating  224  that is applied to the structure  102  that remains adhered, in a chemically unaltered form, to the structure  102  until the aircraft  104  is airborne. As the aircraft  104  travels, the coating  224  sublimes directly into a HMW gas  106 . The coating  224  may be covered prior to use of the aircraft  104  to ensure that it does not sublime prematurely (prior to use of the aircraft). In such a system  214 , the coating  224  would be made of a liquid/solid version of the HMW gas  106  material. 
     The structure  102  or portion of an aircraft where the gas distributors  100  are attached may be selected depending upon the type and location of sound emanating from the aircraft  104  one desires to attenuate. For certain applications of the present invention, it may be desired to ensure that acoustic waves from the aircraft  104  do not travel upstream of the aircraft  104  to prevent said waves from reaching persons in front of the aircraft  104 . For these applications, it would be preferred to place the gas distributors  100  on the nose and/or leading edges of the aircraft  104 . To specifically attenuate engine noise, the gas distributors  100  may be placed on the engines or on the wings adjacent to the engines. To provide control of the aircraft  104  through altering pressure distribution as described above, gas distributors  100  may be placed on or near the wings, tail section, or nose of the aircraft  104 . 
     A HMW gas  106  has a molecular weight of at least 100. For comparison, air has a molecular weight of 29. It is preferred that the HMW gas  106  has a molecular weight above 120 and it is most preferred that the HMW gas  106  has a molecular weight above 140. Many monatomic and polyatomic gases and gas mixtures may be employed as HMW gases  106  to be used in the present invention. Some characteristics favorable to integration with the invention include being colorless, odorless, non-toxic, nonflammable, chemically inert, and thermally stable. Preferred candidate gases should also be readily available from commercial sources and relatively inexpensive. 
     Examples of such HMW gases  106  are monatomic Xenon, Xe, (molecular weight of 131) and polyatomic n-perfluorobutane, C 4 F 10 , (molecular weight of 238). Other refrigerant gases, such as R134A (molecular weight of 102) and R125 (molecular weight of 125), may also be considered. The use of gas mixtures may be especially interesting in terms of the capability of formulating “custom” HMW gases  106  with desired properties. 
     One particular preferred HMW gas  106  for the present invention is sulfur hexafluoride, SF 6 , (molecular weight of 146), which was used for some of the calculations below to further describe the present invention. It possesses many characteristics favorable to integration with the invention, such as being colorless, odorless, non-toxic, nonflammable, chemically inert, and thermally stable. SF 6  is readily available from commercial sources and is inexpensive. SF 6  is commonly used as an insulating gas in electrical equipment and as an etchant in the semiconductor industry. It is chemically inert and stable in the presence of most materials to temperatures of about 500° C. (932° F.). At atmospheric pressure, SF 6  sublimes directly from a solid to a gas. 
     It is also preferred that the HMW gas  106  be provided at a temperature below ambient. As used herein, ambient temperature is the temperature of air outside of an aircraft  104 . As shown further below, the reduction in the speed of sound is increased as the temperature of the HMW gas  106  decreases below ambient. 
     The following briefly describes the mathematical basis of the present invention. 
     The local speed of sound, a, is defined in Equation 1, where γ is the ratio of specific heats, R air  is the specific gas constant for air, and T is the local temperature.
 
 a =√{square root over (γ R   air   T )}  (EQN. 1)
 
     The local Mach number, M, is related to the local velocity, V, and speed of sound using Equation 2. 
     
       
         
           
             
               
                 
                   M 
                   = 
                   
                     V 
                     a 
                   
                 
               
               
                 
                   ( 
                   
                     EQN 
                     . 
                     
                         
                     
                     ⁢ 
                     2 
                   
                   ) 
                 
               
             
           
         
       
     
     The specific gas constant can be related to the universal gas constant, R universal , and the molecular weight, MW, using Equation 3. 
     
       
         
           
             
               
                 
                   
                     R 
                     air 
                   
                   = 
                   
                     
                       R 
                       universal 
                     
                     MW 
                   
                 
               
               
                 
                   ( 
                   
                     EQN 
                     . 
                     
                         
                     
                     ⁢ 
                     3 
                   
                   ) 
                 
               
             
           
         
       
     
     Substituting Equation 3 into Equation 1, the speed of sound is given by Equation 4. 
     
       
         
           
             
               
                 
                   a 
                   = 
                   
                     
                       γ 
                       ⁢ 
                       
                         
                           R 
                           universal 
                         
                         MW 
                       
                       ⁢ 
                       T 
                     
                   
                 
               
               
                 
                   ( 
                   
                     EQN 
                     . 
                     
                         
                     
                     ⁢ 
                     4 
                   
                   ) 
                 
               
             
           
         
       
     
     For constant values of the ratio of specific heats and the universal gas constant, the speed of sound may be decreased by increasing the molecular weight and decreasing the temperature. By decreasing the local speed of sound sufficiently, the local Mach number may reach supersonic values (M&gt;1) in a subsonic velocity flow. 
     The following numerical example demonstrates the desired effect of the present invention by employing the above principles and equations. 
     Assuming an aircraft is flying at an altitude of 6,000 meters (19,685 feet) and a velocity of 130 meters/second (426.5 ft/sec, 252.7 knots). Assuming γ air =1.4 and standard temperature at 6,000 meters altitude of T air =249.2 K (−11.11° F.), the speed of sound is given by Equation 1. 
     
       
         
           
             a 
             = 
             
               
                 
                   
                     γ 
                     air 
                   
                   ⁢ 
                   
                     R 
                     air 
                   
                   ⁢ 
                   T 
                 
               
               = 
               
                 
                   
                     
                       ( 
                       1.4 
                       ) 
                     
                     ⁢ 
                     
                       ( 
                       
                         287 
                         ⁢ 
                         
                           J 
                           
                             kg 
                             - 
                             K 
                           
                         
                       
                       ) 
                     
                     ⁢ 
                     
                       ( 
                       
                         249.2 
                         ⁢ 
                         K 
                       
                       ) 
                     
                   
                 
                 = 
                 
                   316.4 
                   ⁢ 
                   m 
                   ⁢ 
                   
                     / 
                   
                   ⁢ 
                   
                     sec 
                     ⁡ 
                     
                       ( 
                       
                         1 
                         , 
                         
                           038 
                           ⁢ 
                           ft 
                           ⁢ 
                           
                             / 
                           
                           ⁢ 
                           sec 
                         
                       
                       ) 
                     
                   
                 
               
             
           
         
       
     
     The Mach number is given by Equation 2. 
     
       
         
           
             M 
             = 
             
               
                 V 
                 a 
               
               = 
               
                 
                   
                     130 
                     ⁢ 
                     m 
                     ⁢ 
                     
                       / 
                     
                     ⁢ 
                     sec 
                   
                   
                     316.4 
                     ⁢ 
                     m 
                     ⁢ 
                     
                       / 
                     
                     ⁢ 
                     sec 
                   
                 
                 = 
                 0.411 
               
             
           
         
       
     
     At these conditions, the aircraft is subsonic. Assuming sulfur hexafluoride, SF 6 , (γ SF6 =1.095, MW SF6 32 146) is injected into the flow at the same temperature as the freestream air, the local speed of sound is given by Equation 4. 
     
       
         
           
             a 
             = 
             
               
                 
                   
                     γ 
                     SF6 
                   
                   ⁢ 
                   
                     
                       R 
                       universal 
                     
                     
                       MW 
                       SF6 
                     
                   
                   ⁢ 
                   T 
                 
               
               = 
               
                 
                   
                     
                       ( 
                       1.095 
                       ) 
                     
                     ⁢ 
                     
                       ( 
                       
                         
                           8314 
                           ⁢ 
                           
                             J 
                             
                               
                                 kg 
                                 ⁢ 
                                 
                                     
                                 
                                 ⁢ 
                                 mol 
                               
                               - 
                               K 
                             
                           
                         
                         146 
                       
                       ) 
                     
                     ⁢ 
                     
                       ( 
                       
                         249.2 
                         ⁢ 
                         K 
                       
                       ) 
                     
                   
                 
                 = 
                 
                   124.7 
                   ⁢ 
                   m 
                   ⁢ 
                   
                     / 
                   
                   ⁢ 
                   sec 
                 
               
             
           
         
       
     
     The local Mach number is now slightly supersonic. 
     
       
         
           
             M 
             = 
             
               
                 V 
                 a 
               
               = 
               
                 
                   
                     130 
                     ⁢ 
                     m 
                     ⁢ 
                     
                       / 
                     
                     ⁢ 
                     sec 
                   
                   
                     124.7 
                     ⁢ 
                     m 
                     ⁢ 
                     
                       / 
                     
                     ⁢ 
                     sec 
                   
                 
                 = 
                 1.04 
               
             
           
         
       
     
     The effect of temperature can be seen if the sulfur hexafluoride is injected at its sublimation temperature of −63.9° C. (209.25 K, −83.0° F.). The speed of sound is decreased by about 8.4%. 
     
       
         
           
             a 
             = 
             
               
                 
                   
                     γ 
                     SF6 
                   
                   ⁢ 
                   
                     
                       R 
                       universal 
                     
                     
                       MW 
                       SF6 
                     
                   
                   ⁢ 
                   T 
                 
               
               = 
               
                 
                   
                     
                       ( 
                       1.095 
                       ) 
                     
                     ⁢ 
                     
                       ( 
                       
                         
                           8314 
                           ⁢ 
                           
                             J 
                             
                               
                                 kg 
                                 ⁢ 
                                 
                                     
                                 
                                 ⁢ 
                                 mol 
                               
                               - 
                               K 
                             
                           
                         
                         146 
                       
                       ) 
                     
                     ⁢ 
                     
                       ( 
                       
                         209.25 
                         ⁢ 
                         K 
                       
                       ) 
                     
                   
                 
                 = 
                 
                   114.2 
                   ⁢ 
                   m 
                   ⁢ 
                   
                     / 
                   
                   ⁢ 
                   sec 
                 
               
             
           
         
       
     
     The Mach number is increased by about 9.6%. 
     
       
         
           
             M 
             = 
             
               
                 V 
                 a 
               
               = 
               
                 
                   
                     130 
                     ⁢ 
                     m 
                     ⁢ 
                     
                       / 
                     
                     ⁢ 
                     sec 
                   
                   
                     114.2 
                     ⁢ 
                     m 
                     ⁢ 
                     
                       / 
                     
                     ⁢ 
                     sec 
                   
                 
                 = 
                 1.14 
               
             
           
         
       
     
     The table below summarizes the results of this numerical example. 
     Altitude=6,000 m (19,685 ft) 
     Velocity=130 m/sec (426.5 ft/sec) 
     
       
         
               
               
               
               
             
               
               
               
               
               
             
           
               
                   
                   
               
               
                   
                   
                 SF 6  Injection, 
                 SF 6  Injection, 
               
               
                   
                 No SF 6   
                 Ambient 
                 Low 
               
               
                   
                 Injection 
                 Temperature 
                 Temperature 
               
               
                   
                   
               
             
             
               
                   
               
             
          
           
               
                   
                 SF 6  Temperature 
                 None 
                 249.2 
                 209.3 
               
               
                   
                 ° K (° F.) 
                   
                 (−11.11) 
                 (−83.0) 
               
               
                   
                 Speed of Sound, a 
                 316.2 
                 124.7 
                 114.2 
               
               
                   
                 m/sec (ft/sec) 
                 (1,038) 
                 (409.1) 
                 (374.7) 
               
               
                   
                 Mach number, M 
                 0.411 
                 1.04 
                 1.14 
               
               
                   
                   
               
             
          
         
       
     
     Referring to  FIGS. 4-6 , representative results of computational fluid dynamics (CFD) analyses of the invention are depicted. The CFD analysis is a two-dimensional, viscous, non-reacting, unsteady numerical solution. 
       FIG. 4  shows a computational fluid dynamic analysis model. Air enters the inflow plane on the left of the figure at a subsonic Mach number of about 0.78. SF 6  gas is injected transversely into the air flow at the lower boundary, downstream of the inflow. 
       FIG. 5  shows the SF 6  mass fraction distribution for injection of SF 6  gas transversely into the subsonic air flow. The injection distributes a pocket of HMW gas into the main subsonic flow. 
       FIG. 6  shows the Mach number contour map with SF 6  injection. The subsonic flow enters the pocket of high molecular weight SF 6  gas, where the speed of sound is substantially reduced. The flow becomes locally supersonic in the SF 6  gas pocket and reaches a supersonic Mach number of about 1.1. Disturbances produced downstream of the supersonic flow region cannot propagate upstream. 
     What is described are specific examples of many possible variations on the same invention and are not intended in a limiting sense. The claimed invention can be practiced using other variations not specifically described above.