Abstract:
An injector for use with the rocket thruster has a plurality of fuel ports separated from a plurality of oxidizer ports. The oxidizer and fuel ports are paired together directing their respective fluids along a path with radial and tangential components so that the two fluids impinge at a predetermined spaced apart distance from the chamber wall of the combustion chamber at an impingement track. By providing the fuel at a steeper angle relative to the chamber walls than the oxidizer, the fuel can be utilized to provide a fuel rich zone near the chamber walls to assist in cooling the chamber walls during operation.

Description:
ORIGIN OF THE INVENTION 
   The invention described herein was made by employees of the United States Government and may be manufactured and used by or for the Government for Governmental purposes without the payment of any royalties thereon or therefore. 

   BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   This invention relates to a liquid propellant injector for use with a rocket thruster, more particularly to an improved injector defining combustion chamber therein. 
   2. Prior Art 
   Various injectors have been utilized with rocket propulsion systems in the past. U.S. Pat. No. 5,765,361 shows an injector which provides the oxidizer as well as the fuel axially in the direction of a nozzle. The igniter engine shown in FIG. 1A initially provides oxidizer to start the combustion process. Axially directed propellants typically take a relatively long combustion chamber to ensure that propellants mix and completely combust prior to exiting the nozzle. 
   U.S. Pat. No. 4,586,226 operates somewhat similarly to the &#39;361 patent. Fuel is provided axially in the direction of the nozzle in the illustrated injector from a face plate at the chamber head end. In this design, the oxidizer is directed at an angle to intersect the flow path of the fuel and then proceed axially towards the nozzle (not shown). 
   In designs which axially direct propellants, the combustion chamber walls are typically cooled by film cooling or some other means, to prevent material degradation. These cooling concerns are believed to create a loss in performance. Additionally, traditional injectors often require a long time to fabricate and are usually costly to produce. 
   In order to overcome potential disadvantages of a long combustion chamber, numerous efforts have been undertaken to provide tangentially directed propellants. In U.S. Pat. No. 3,937,012, the oxidizer is initially supplied axially and inwardly to the fuel which then impinges upon a impeller to provide a tangential aspect to the flow of propellants. The oxidizer provides a cooling veil over the top surface of the impeller. (Column  4 , lines  40 - 45 .) The impeller design is believed to add to the complexity of the injector and possibly provide a point of failure and increased costs. 
   U.S. Pat. No. 3,640,072 shows a simpler combustion chamber for use with the rocket that directs propellants tangentially into the combustion chamber to the combustion sidewalls rather than actually at the head end of the chamber. The liquid propellants impinge along the chamber walls which is believed to impose the greatest thermal loads at the point of impingement along the walls. These thermal stresses are believed to be high enough to drastically reduce the life of the components. 
   While there have been a number of improvements in basic combustion chamber design, a need still exists to provide a low cost injector which is robust, requires little maintenance and may be reused. A need also exists to provide an engine combustion chamber having reduced complexity, cost of fabrication and overall weight. 
   SUMMARY OF THE INVENTION 
   Consequently, it is a primary object of the present invention to provide a low cost injector to be used in rocket engines or other combustion chambers. 
   Another object of the invention is to direct fuel and oxidizer within the combustion chamber so that the flow vortices keep low temperature fluids on the wall while the hot gas is at the center core. 
   Another object of the present invention is to provide an impingement trace internal to the combustion chamber walls with the fuel directed closer to the walls than the oxidizer. 
   Accordingly, a tracing impingement injector has a plurality of fuel ports separated from a plurality of oxidizer ports. The oxidizer and fuel ports direct fuel and oxidizer into the combustion chamber with a radial and tangential component such that the two fluids impinge a predetermined spaced apart distance from the chamber wall. It is further preferred that the fuel be introduced closer to the chamber walls so that the fuel provides a fuel rich zone near the wall to assist in cooling to the wall and to eliminate the wall oxidation during operation. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The particular features and advantages of the invention as well as other objects will become apparent from the following description taken in connection with the accompanying drawings to which: 
       FIG. 1  is a top plan view of the tracing impingement injector; 
       FIG. 2  is a cross-sectional view taken along the line A—A of  FIG. 1 ; and 
       FIG. 3  shows the injector installed relative to a downstream nozzle in a rocket thruster. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
   Referring to  FIGS. 1 and 2 , a liquid propellant injector  10  is disclosed for use in a rocket thruster  12  as shown in FIG.  3 . Referring back to  FIGS. 1 and 2 , the injector  10  is comprised of a housing  14  containing a combustion chamber  16  therein. The combustion chamber is defined by chamber walls  18  which surround the chamber  16 . The chamber walls  18  have ports  20 , 22  which will be explained in more detail below which provide fuel and oxidizer to the combustion chamber  16 . Fuel and oxidizer are often referred to as propellants herein. By fabricating the housing  14  with a three-dimensional computer aided design model, a laser engineered net shaping (LENS) machine was able to shape the combustion chamber  16  in the housing  14  and provide ports  20 , 22  as illustrated. Ports  20  are preferably utilized to provide oxidizer into the combustion chamber. Ports  22  are preferably utilized to provide fuel into the combustion chamber  16 . 
   Inlet ports  24  receive oxidizers such as liquid oxygen from a supply. Inlet ports  26  receive fuel from a fuel supply. From the inlet ports  24 , 26 , the respective propellant fluid, whether fuel or oxidizer, is directed through respective manifolds  28 , 30 . Oxidizer manifold  28  is illustrated as a concentric ring along an upper portion of the housing  14  which is in communication with inlet port  24 . Fuel manifold  30  is illustrated in a similar manner in communication with inlet port  26  in FIG.  2 . As can be seen in  FIG. 2 , fuel enters inlet port  26  and passes into manifold  30  through duct  32 . Once in the manifold  30 , the fuel can be directed about the manifold  30  and dispensed through flues  34  into outlet ports  22 . Once in the outlet ports  22 , the fuel can then pass into the combustion chamber  16  as shown in FIG.  1 . 
   As illustrated in  FIG. 1 , there are fourteen outlet ports  22  for fuel to be provided to the combustion chamber  16 . There are also fourteen outlet ports  20  for oxidizers to be provided into the combustion chamber  16 . The manifold/inlet/outlet port arrangement for providing the oxidizer in the preferred embodiment is very similar to the arrangement for the fuel. The oxidizer is provided from the inlet port  24  to the manifold  28  and then through the outlet ports  20  into the combustion chamber  16 . 
   As shown in  FIG. 1 , the fuel outlet ports  22  introduce fuel closer to the chamber walls  18  than the oxidizer is introduced. This is the result of the geometry of the construction of the outlet ports  20 , 22 . This feature is not only believed to assist in cooling the injector  10 , but also to provide a fuel rich environment at the rear wall region to avoid the wall oxidation. 
   As illustrated in  FIG. 1 , the fuel and oxidizer injected from the outlet ports  20 , 22  impinges at a predetermined impingement track  36  which is spaced a predetermined distance from the chamber walls  18 . Specifically in the preferred embodiment, the combustion chamber  16  has an internal diameter of 2.35 inches while the interior diameter of the impingement track has a diameter of 1.95 inches which is 0.20 inches spaced from the chamber walls  16 . The spacing of the impingement track  36  from the chamber wall  18  has been found to assist in providing a fuel rich zone near the chamber wall  18  which assist in providing cooling to the chamber wall  18 . 
   The ports  20 , 22  are preferably aligned in a port plane  38  illustrated in  FIG. 2  which happens to coincide with the impingement track  36  which lies in an impingement plane. As can be seen in  FIG. 2 , the size of the oxidizer ports  20  are slightly greater than that of the fuel ports  22 . This has been found to assist in an ability to provide fuel such as RP-1, at 1.15 lbm/sec while providing liquid oxidizer, at 2.96 lbm/sec. However, by varying the diameters and pressures at which the fuel and oxidizers supplied to the combustion chamber  16 , various flow rates can be achieved. For instance, the above flow rates have been achieved when supplied both fuel and oxidizer at 1200 psia. Other flow rates may be desired for other propellants or other performance criteria. 
   The ports  20 , 22  may be angled at other orientations other than the orientations provided in FIG.  1 . In  FIG. 1 , the oxidizer ports  20  are angled at about 40 degrees from a radius extending from a center axis  21  of the combustion chamber  16  extending through the any particular port  20 . The fuel port  22  is angled at about 20 degrees relative to a radian extending through the center axis  21  of the combustion chamber  16  and the respective fuel port  22 . By varying these angular relationships, the impingement track  36  may be moved closer or farther away from the chamber wall  18 . Although the impingement track  36  is illustrated as being circular, it could take a number of forms depending on the shape of the chamber walls  18  and the orientation of the ports  20 , 22  as well as the relative size of the ports and pressures at which the propellants are introduced into the combustion chamber  16 . 
   Seal grooves  40  provide a location for seals to extend circumferentially around the combustion chamber  16  to assist in preventing hot gasses in the combustion chamber  16  from mixing with either of the propellants in either of the manifolds  28 , 30 . With some embodiments, final machining may be necessary to incorporate the fuel and oxidant inlet ports  24 , 26  and/or provide surface finishes for sealing with other components. 
   As the fuel injects into the combustion chamber  16 , it is preferred that the fuel is introduced closer to the wall  18 . As the propellants mix, a fuel rich zone near the wall  18  provides cooling. The propellants then migrate toward the head end  42  shown in FIG.  3 . At the head end  42 , the combustion products flow inward forming a circulating core or vortex and expand as they flow down the core toward the nozzle  44  past the throat. 
     FIG. 3  shows the chamber head housing  46  which is connected such as by eight bolts, to the injector  10  and the structural jacket  48 . The chamber head housing  46  has an interior passage  50  which receives an igniter (not shown) therein. The igniter transmits a flame through an orifice  52  into the combustion chamber  16 . 
   Below the injector  10  is the structural jacket  48  which surrounds the nozzle insert  54 . A number of cooling ducts  56  extend around the circumference of the nozzle insert  54  and may be cooled by a number of fluids, including the fuel utilized as a propellant. The nozzle insert  54  has been constructed with oxygen free high conductivity copper, however, other appropriate materials may also be utilized. The structural jacket  48  has been constructed of 304L stainless steel while the other parts of the rocket  12  have been constructed of Inconel  718  or other appropriate material. 
   In order to assist cooling of the head end  42  of the injector and the lower portion of the chamber head housing  46 , a Rigimesh™ plate  58  may be utilized which has a plurality of microscopic openings therethrough which has been found to assist in maintaining the chamber head housing  46  at acceptable temperatures during the use of the rocket  12 . 
   Numerous alternations of the structure herein disclosed will suggest themselves to those skilled in the art. However, it is to be understood that the present disclosure relates to the preferred embodiment of the invention which is for purposes of illustration only and not to be construed as a limitation of the invention. All such modifications which do not depart from the spirit of the invention are intended to be included within the scope of the appended claims.