Abstract:
Improved rocket nozzle designs for vehicles with nozzles embedded in or protruding from surfaces remote from the desired thrust axis. The nozzle configurations are for rocket vehicles where the nozzles are not located at the optimal thrust axis of the vehicle. Two examples include nozzles located on the forward end of the vehicle (also called tractor nozzles) and attitude control nozzles located on the periphery of the vehicle. More particularly, the disclosed nozzle shapes enhance the axial thrusts and/or maneuver torques on the vehicle. These unconventional nozzle shapes improve vehicle performance.

Description:
CROSS REFERENCE TO RELATED APPLICATION(S) 
       [0001]    N.A. 
       U.S. GOVERNMENT RIGHTS 
       [0002]    N.A. 
       BACKGROUND 
       [0003]    1. Field 
         [0004]    Disclosed are nozzle configurations for rocket vehicles where the nozzles are not located at the optimal thrust axis of the vehicle. Two examples include nozzles located on the forward end of the vehicle (also called tractor nozzles) and attitude control nozzles located on the periphery of the vehicle. More particularly, the disclosed nozzle shapes enhance the axial thrusts and/or maneuver torques on the vehicle. 
         [0005]    2. Description of the Related Art 
         [0006]    In conventional rocket nozzle design, the flow and pressure in the nozzle is evaluated as a simple function of the local area to the throat area.  FIG. 1  illustrates an axisymmetric nozzle  10  as known from the prior art. The average pressure at each station in the nozzle  10  can be found by inverting the following equation relating the area ratio A/A t , pressure ratio P c /P and specific heat ratio gamma, 7. 
         [0000]    
       
         
           
             
               
                 
                   
                     A 
                     At 
                   
                   = 
                   
                     
                       
                         ( 
                         
                           
                             P 
                             c 
                           
                           / 
                           P 
                         
                         ) 
                       
                       
                         
                           γ 
                           + 
                           1 
                         
                         
                           2 
                            
                           γ 
                         
                       
                     
                     
                       
                         
                           [ 
                           
                             
                               γ 
                               + 
                               1 
                             
                             2 
                           
                           ] 
                         
                         
                           
                             γ 
                             + 
                             1 
                           
                           
                             2 
                              
                             
                               ( 
                               
                                 γ 
                                 - 
                                 1 
                               
                               ) 
                             
                           
                         
                       
                        
                       
                         
                           
                             2 
                             
                               γ 
                               - 
                               1 
                             
                           
                            
                           
                             [ 
                             
                               
                                 
                                   ( 
                                   
                                     
                                       P 
                                       c 
                                     
                                     / 
                                     P 
                                   
                                   ) 
                                 
                                 
                                   
                                     γ 
                                     - 
                                     1 
                                   
                                   γ 
                                 
                               
                               - 
                               1 
                             
                             ] 
                           
                         
                       
                     
                   
                 
               
               
                 
                   ( 
                   1 
                   ) 
                 
               
             
           
         
       
     
         [0000]    A is local flow area measured in square inches;
 
At is the sonic throat area measured in square inches;
 
P c  is the chamber total pressure measured in pounds per square inch;
 
P is local flow pressure measured in pounds per square inch; and
 
γ is specific heat ratio (unit-less).
 
         [0007]    Even this calculation has its limits in that the pressure, p, is the average at a station and the wall pressure may be more or less than the average p, a function of P c  &amp; (A/A t ). This technique has been successfully employed to design axisymmetric nozzles following simple rules. Gas flows into the nozzle  10  at subsonic speed and through a converging portion  12  that terminates at throat  14 . The gas then flows at supersonic speed through diverging portion  15  and through exit  16 . The supersonic flow in the nozzle  10  responds to changes in the nozzle wall contour  17  through expansion or shock waves. These waves will travel from their origin across the nozzle  10  to the opposite side where they reflect and cross back and forth as the flow accelerates from the throat  14  to the exit  16 . These waves are called characteristics or Mach waves and in the days before Computational Fluid Dynamics (CFD), were used to design and analyze the flow in a nozzle  10 . 
         [0008]      FIG. 2  shows how these Mach waves  18  influence the flow and pressure in the diverging portion  15  of a simplified two dimensional nozzle with a supersonic starting Mach number. This nozzle has a single wall angle change, β, from an axial throat  14  to a 15° expansion. The flow in the nozzle can be broken into several different zones of uniform but strikingly different properties. Zone  0  is the supersonic flow at the throat  14  of the nozzle. An exemplary Mach number is 1.5 and the pressure is 29% of the total pressure of the approaching flow. Zone  1  is downstream of the 15° outward turn. The exemplary Mach number has increased to 1.944 and the pressure has dropped to 14.5%. In Zone  1 , the flow properties are uniform and the flow direction is parallel with the nozzle surface. A number of waves can be seen originating at the transition corner  22  and fanning out. This is the well known Prandtl-Meyer Expansion. These Mach waves  18  expand from the origin due to the increasing Mach number as the flow turns and accelerates around the transition corner  22 . These waves cross the nozzle centerline  24  generating a Zone  2 . In Zone  2 , the crossing of the waves from opposite sides of the nozzle cause more turning and acceleration of the flow. The exemplary flow in Zone  2  is now moving parallel to the axis  24  at Mach 2.41 and 6.5% of the total pressure. As the waves continue to crisscross the nozzle, no more Zones of uniform properties are formed in this example and each point in the flowfield has varying Mach number, pressure and flow direction all resulting from the interactions with these waves. If the nozzle is cut off at a fixed length, the average properties at the exit  16  will determine the resulting thrust. Since the flow in not perfectly uniform and axial, a nozzle efficiency term is usually applied for delivered nozzle performance. 
         [0009]    In a more general case, the wall contour  17  is not uniform as in this example and the pressure acting on the wall is affected directly by changes in wall angle and waves from the opposite side of the nozzle. Since the Mach waves move at the local speed of sound in the gas, they move away from the wall and are swept downstream by the supersonic flow. This means that wall angle changes affect the local pressure and the pressure on the opposite wall well downstream of the origin of that change. If, for example, the lower wall  26  was terminated at point A, the flow would expand out and downward causing another expansion fan to travel from A to B. Since B is beyond the exit of the nozzle the removal of this lower wall section will not change the pressures on the upper wall  28 . 
         [0010]    In most rocket propelled vehicles, one or more rocket nozzles are located at the aft end and aligned close to the vehicle axis to convert most of the thrust in flight direction. Some applications exist that preclude this aft location for nozzles. Exemplary tractor nozzle applications are the TOW missile and the escape motor for the Orion, Crew Exploration Vehicle. In another application, multiple rocket nozzles are arrayed around the vehicle for maneuvering. For both of these applications, using conventional nozzles can have a detrimental effect on vehicle performance. 
         [0011]    Referring to  FIG. 3 , the launch abort system  30  of an Orion Crew Exploration Vehicle  32 , or similar rocket, has a payload  34  that includes a crew module  36  and a service module  38  located forward of booster rocket  40 . 
         [0012]    In the event that it becomes necessary to abort the mission prior to separation of the booster rocket  40  from the payload  34 , abort motor  42  is ignited generating propellant gases that are expelled through nozzles  44 . Nozzles  44  are angularly disposed in an aftward direction to the intended direction of flight of the vehicle and generate a thrust effective to separate the launch abort system and crew module  30  from the remainder of the vehicle  32 . This propulsion system has a number of limitations. Hot propellant gases expelled by the nozzles  44  impact the abort motor  42  and crew module  36 . These components must be designed to withstand the high temperatures that may be generated by impingement of the propellant gases. One approach, that was utilized on the Apollo program, is to cant the nozzles  44  at an extreme angle to expel the hot propellant gases sufficiently outward from the rocket components to avoid the most severe temperature increases. However, deviation of the nozzle direction from directly aftward causes a loss of thrust requiring abort motor  42  to be charged with additional propellant. 
         [0013]    These limitations are illustrated in more detail in  FIGS. 4 and 5 . Referring to  FIG. 4 , when the nozzle  44  is canted at an angle, α, relative to direction of flight  26 , the cant angle on the multiple nozzles causes a loss in thrust that is roughly equal to (1−cos α) which at 30 degrees of cant represents a thrust loss of 13%. 
         [0014]    Nozzles  44  project outward from exterior walls  48  of the abort motor  42  that arrayed around motor centerline  47 . The abort motor  42  includes a propellant  49  that when ignited generates propellant gases. The propellant gases pass through nozzle throat  14  and are accelerated through divergent portion  15  of the nozzle. If the cant angle, α, of the nozzle is greater than the divergence angle, β, the wall  50  of the nozzle  44  generates a negative thrust due to the local nozzle pressures acting in a rearward direction. The opposite effect is seen on wall  53 . Here the wall pressures  55  have a larger forward projected area and contribute more to the net thrust of the nozzle  44 . However, with the conventional nozzle  44 , the thrust is significantly less than an axially directed nozzle (due to the cosine loss from the previous paragraph). 
         [0015]    Referring now to  FIG. 5 , the abort motor  42  must function from sea level static to very high altitudes. The most important condition is likely to be at the maximum dynamic pressure (max Q) point in the assent trajectory. This is the point of highest drag and stress on the vehicle and is the most challenging for the design of the escape system. Under these conditions, the approaching supersonic flow  52  causes a bow shock wave  54  and elevated pressures that deflect the rocket plume  56  back towards the motor exterior walls  48  and crew module  36 . With reference to  FIG. 6 , this results in impingement of underexpanded rocket plume  56  on the exterior wall  48  at attachment point  58 . The plume impingement separates the boundary layer  60  which on reattachment, results in a hot spot that may degrade the exterior wall  48  structural capability. 
         [0016]    Referring back to  FIG. 5 , making the situation worse is the fact that the attachment point  58 ′,  58 ″,  58 ′″ shifts with flight speed. A second problem is that despite a large cant angle, α, the exhaust plume  56 ′ still interacts with the crew module  36 . Increasing the distance between the nozzles  44  and crew module  36  will reduce the heating caused by the plume  56 ′ but will not avoid the flow reaching the forward surface of the crew module. 
         [0017]    The crew module  36  typically includes roll control nozzles  62  that control module roll while in orbit. As shown in  FIG. 7 , the roll control nozzle  62  has an exit portion that is flush with the perimeter  64  of the crew module  36  or other space vehicle. The function of this rocket thruster is to control the orientation of the vehicle by providing a torque about its center of mass (CG). A conventional roll control nozzle  62  is installed with one exit side  65  flush with the vehicle perimeter  64 , the nozzle is then scarfed so that the opposing exit side  67  is also flush with the vehicle perimeter (no portion of the nozzle protrudes from the vehicle). The expanding gases in the nozzle apply a force F 1  along the nozzle axis and also F 2  normal to the nozzle axis. The intended roll controlling torque is F 1 *R 1  while the scafing creates a counter torque (F 2 *R 2 ) on the unbalanced surface A 2 . The net torquer, applied to the vehicle is then: 
         [0000]      τ= F 1 *R 1 −F 2 *R 2  (2) 
         [0000]    Compensation for the counter torque requires more propellant to be consumed for a given maneuver than desired. 
         [0018]    There remains a need for improved rocket nozzles, such as for a launch abort system, or the capsule attitude control system, that overcomes this loss of vehicle performance. 
       BRIEF SUMMARY OF THE INVENTION 
       [0019]    Disclosed are nozzle configurations for rocket vehicles where the nozzles are not located at the optimal thrust axis of the vehicle. Two examples include nozzles located on the forward end of the vehicle (also called tractor nozzles) and attitude control nozzles located on the periphery of the vehicle. More particularly, the disclosed nozzle shapes enhance the axial thrusts and/or maneuver torques on the vehicle 
         [0020]    The details of one or more embodiments of the invention are set forth in the accompanying drawings and the description below. Other features, objects and advantages of the invention will be apparent from the description and drawings, and from the claims. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0021]      FIG. 1  illustrates an axisymmetric rocket nozzle as known from the prior art. 
           [0022]      FIG. 2  illustrates characteristic waves of a propellant gas expanded in the prior art axisymmetric rocket nozzle of  FIG. 1 . 
           [0023]      FIG. 3  illustrates in a launch abort system with tractor nozzles as known from the prior art. 
           [0024]      FIG. 4  illustrates a loss of thrust associated with the prior art tractor nozzles of  FIG. 3 . 
           [0025]      FIG. 5  illustrates plume impingement occurring with the prior art tractor nozzles of  FIG. 3 . 
           [0026]      FIG. 6  illustrates separation of the boundary layer and formation of a hot spot as a result of the plume impingement of  FIG. 5 . 
           [0027]      FIG. 7  illustrates a roll control nozzle as known from the prior art. 
           [0028]      FIG. 8  illustrates a first embodiment of tractor nozzle that does not have the limitations of the prior art. 
           [0029]      FIG. 9  illustrates a second embodiment of tractor nozzle that does not have the limitations of the prior art. 
           [0030]      FIG. 10  is a front planar view of support web for the nozzles disclosed herein. 
           [0031]      FIG. 11  illustrates a third embodiment of tractor nozzle that does not have the limitations of the prior art. 
           [0032]      FIG. 12  illustrates a first embodiment of a nozzle for a roll control motor that does not have the limitations of the prior art. 
           [0033]      FIG. 13  illustrates a second embodiment of a nozzle for a roll control motor that does not have the limitations of the prior art. 
           [0034]      FIG. 14  illustrates a third embodiment of a nozzle for a roll control motor that does not have the limitations of the prior art. 
       
    
    
       [0035]    Like reference numbers and designations in the various drawings indicated like elements. 
       DETAILED DESCRIPTION 
       [0036]      FIG. 8  illustrates a first embodiment of a nozzle  102  that does not have the disadvantages described hereinabove. The nozzle  102  protrudes from an exterior wall  48  of a rocket motor, such as abort motor  42 . The nozzle  102  protrudes from the exterior wall  48  and is angularly disposed in an aftward direction relative to the direction of flight  46 . An optional faring  85  can be used to reduce the drag on the projecting nozzle. The nozzle  102  has an aft surface portion  104  that is adjacent to the exterior wall  48  and a fore surface  106 . When viewed relative to the direction of flight  46 , aft surface portion  104  extends rearward for a lesser distance than the opposing fore surface  106 . Nozzle  102  is essentially a fragment of the nozzle  44  ( FIG. 4 ). Broken line  76  identifies in phantom that portion of the nozzle  44  omitted from the nozzle  102 . More precisely, that portion of the aft surface  104  of the nozzle to the rear of the point of attachment of the last characteristic (Mach wave  78 ) that can reach the fore surface  106  is removed. This will increase the thrust by 10% or more because some of the negative thrust generated by the pressure acting on the aft surface  104  is eliminated. 
         [0037]    The rocket plume  56  reacts to the change in nozzle  102  contour by turning and accelerating to a higher Mach number. This thrust increase is due to both the greater expansion and more axial jet angle, that is the exhaust flow is nearly axial in direction. By nearly axial, it is meant that axial jet angle has a deviation of less than 10° relative to the longitudinal axis  107  of the vehicle. By omitting the forward facing wall of the aft portion  104  of the nozzle, the flow is allowed to turn onto the exterior wall  48  of the motor case. This results in the flow remaining attached with little or no boundary layer separation keeping the heat flux similar to that within the nozzle  102 . This thermal environment is then more benign and consistent over the flight speeds. 
         [0038]      FIG. 9  illustrates in cross-sectional representation a nozzle  110  in accordance with a second embodiment that does not have the disadvantages identified hereinabove. In this embodiment, propellant gas  112  generated by burning propellant  49  passes through subsonic converging portion  114  that is supported by structural web  116 , that is shown in end view in  FIG. 10 . 
         [0039]    Returning to  FIG. 9 , the propellant gases  112  then flow through an internal throat  118  while traveling at an angle that is approximately 90° from the motor axis  47 . The forward and aft surfaces of the nozzle  110  are connected by a series of webs ( 120 , see  FIG. 10 ) that interrupt the throat  118  to carry the loads. The flow is first turned through a large angle by Prandtl-Meyer expansion waves  94 . A forward surface  122  of the nozzle  110  is defined by the expansion waves  94 . The forward surface  122  is exposed to a high but declining pressure. The rear surface  124  is exposed to a uniform and low pressure. As the flow exits the outer diameter of the motor, a small rearward deflector  126  in the projected forward surface  122  generates a local high pressure increasing the thrust and making the emerging flow  56  more axial. If the resulting shock wave  57  does not reach the aft surface  124 , no negative thrust is created. 
         [0040]    As shown in  FIG. 11 , if a fully submerged nozzle  130  is desired, that is nothing sticking out into the airstream; the diverging expansion portion  132  is terminated at the motor case diameter  48 . An increase in thrust can be achieved by adding a slight rearward deflector  126  to the end of the forward nozzle surface  122 . As in the preceding embodiment, deflector  126  is located so that the resulting shock wave  54  does not reach the aft surface  124 . 
         [0041]    For vehicle maneuvering thrusters, the counter torque of prior art integrations exemplified by  FIG. 7  are mitigated by embodiments illustrated in  FIGS. 12-14 . In a first embodiment illustrated in  FIG. 12 , the surface A 2 ′ that generates the counter torque, is displaced away from the centerline  24  of the nozzle  140 , in a direction towards the vehicle center, CG. This reduces the pressure, P′, acting on the surface A 2 ′ and reduces its moment arm R 2 . The torque, τ′ is increased because: 
         [0000]      τ′= F 1* R 1 −F 2′* R 2′  (4) 
         [0000]        F 2′= P′*A 2′&lt; P*A 2  (5) 
         [0000]    P′ is much less than P and R 2 ′ is less than R 2  from the  FIG. 7  embodiment. 
         [0042]    In a second embodiment for a roll control nozzle  142 , shown in  FIG. 13 , the favorable surface of the diverging expansion portion  132  is elongated by extension portion A 3  thereby further increasing the roll torque, τ″. In this embodiment, 
         [0000]      τ″=( F 1* R 1+ F 3* R 3)−( F 2* R 2)  (6) 
         [0000]    and the contribution to desired torque is enhanced by (F 3 *R 3 ) while the undesired counter-torque is reduced due to a small value for F 2  as disclosed in the preceding embodiment. 
         [0043]    In a third embodiment for a roll control nozzle  144 , shown in  FIG. 14 , the nozzle maximizes torque production. In this embodiment the flow is first rapidly expanded by a centered Prandtl-Meyer Expansion  146 . This expansion can turn the flow by 90° or more while lowering the pressure to a small fraction of its initial value. The exterior most portion  148  of the nozzle  144  is the over-turned, that is turned faster than the Prandtl-Meyer Expansion would create, thus causing the local nozzle wall pressure to rise due to a net compression  150  of the flow. This causes a further deflection of the flow shown by the jet  152 . This increases the torque due largely to the change in the thrust vector direction and its larger resulting moment arm R 1 ′″ even if the thrust magnitude F 1 ′″ is unchanged. 
         [0044]    One or more embodiments of the present invention have been described. Nevertheless, it will be understood that various modifications may be made without departing from the spirit and scope of the invention. Accordingly, other embodiments are within the scope of the following claims.