US3286474A - Hoop segmented injector and combustor - Google Patents

Description

Nov. 22, 1966 H. A. HOCHE, JR.. ET AL 3,

HOOP SEGMENTED INJECTOR AND COMBUSTOR 4 Sheets-Sheet 1 Filed Dec. 5, 1962 INVENTORS HERBERT A. HOCHE JR. WILLIAM R. SEUGLING ATTORN EY Nov. 22, 1966 HOCHE, JR, ET AL 3,285,474

HOOP SEGMENTED INJECTOR AND COMBUSTOR Filed Dec. 5, 1962 4 Sheets-Sheet 2 27 2a 30G 22 2o 24 25 l 29 k 2a\ M INVENTORS HERBERT A. HOCHE JR, By WILLIAM R. SEUGLING A'ITORNEY 1966 H. A. HOCHE, JR.. ET AL 3,

HOOP SEGMENTED INJECTOR AND COMBUSTOR Filed Dec. 5, 1962 4 Sheets-Sheet 5 INVENTORS HERBERT A. HOCHE JR. BY WILLIAM R. SEUGLING ATTORNEY Nov. 22, 1966 A, HOCHE, JR.. ET AL 3,286,474

HOOP SEGMENTED INJECTOR AND COMBUSTOR 4 Sheets-Sheet 4 Filed Dec. 5, 1962 INVENTORS HERBERT A. HOCHE JR.

By WILLIAM R. SEUGLING ATTORNEY United States Patent Oflflce 3,286,474 Patented Nov. 22, 1966 3 286 474 HOOP SEGMENTED Il lJE( 3TOR AND COMBUSTOR Herbert A. Hoche, Jr., Woodland Hills, and William R.

Seugling, Van Nuys, Calif., assignors to North American Aviation Inc.

Filed Dec. 5, 1962, Ser. No. 243,202 7 Claims. (Cl. 60260) This invention relates to injectors and combustion chambers of rocket engines. More particularly, this invention relates to annular combustion chambers.

Rocket engine thrust chambers in the past have comprised a cylindrical combustion chamber, a converging throat area and a diverging area where expansion of the combustion gases occurs. It is common practice to dispose these component in axial alignment which accordingly requires a relatively great length to accomplish the designed function. Approximately 20 to 30 percent only of the total achievable thrust is produced on the diverging nozzle walls, the balance being produced over the combustion chamber area. More recently developed thrust chambers utilize an annular throat construction to obviate some of the disadvantages of axially aligning all of the components. This is disclosed in patent application Serial No. 73,726 filed December 5, 1960, and applications cited therein, all of which are assigned to the assignee of the instant invention. However, even in these types of constructions, it is necessary to make portions of the structure heavy and unwieldly for purposes of strength. It is to obviate some of the disadvantages of these prior art structures that forms the purpose of the invention.

Briefly, the invention comprises hoop segments, usually a plurality, for rocket engines, the hoops together forming a combustion chamber. Each hoop or hoop segment conveys a fiuid therethrough to cool the hoops and to heat the fluid which may be a propellant. The propellant i injected into the combustion chamber and ignited. After combustion the hot gases pass through throat areas formed between hoops, expanded in an expansion chamber and defiected by some means resulting in useful thrust.

It is accordingly an object of this invention to provide for an improved injection and combustion system.

It is a more particular object of this invention to provide for an injection and combustion chamber which utilizes hoop segments which regeneratively cool the combustion chamber while providing chamber strength and light weight for annular type nozzles.

Other and more specific objects of this invention will become apparent from the description and accompanying drawings in which;

FIG. 1 is a cross-sectional view of an expansion-deflection nozzle illustrating a previous solution to problems of rocket design.

FIG. 2 is a perspective view of two hoop segments and injector.

FIG. 3 is a cross-sectional view along the lines 33 of FIG. 2.

in the prior art.

.in much lowered thrust levels.

FIG. 10 is a partial cut-away cross section of another embodiment of this invention.

FIG. 11 is a partial perspective view of another embodiment of this invention.

FIG. 12 is a view taken in the direction of line 1212 of FIG. 11.

FIG. 13 is a view looking in the direction of line 1313 of FIG. 11.

FIG. 14 is a cross-sectional of a further modification of this invention.

FIG. 1 discloses a prior art device described in the aforementioned patent application and is an advance in the art over previously constructed rocket engines in that substantial savings in length and weight has been accomplished through the expansion-deflection principle. As shown in FIG. 1, propellants are injected into combustion chamber 1 and ignited. The gases are then passed through throat area 2 which is annular in configuration where the gases are accelerated. After passing through this throat area 2 the gases are expanded and then deflected by skirt 5. The gases exhaust at the rear of the nozzle (not shown). Skirt 5 is the bell portion of the nozzle. This .type of configuration saves weight and length, thus resulting in a more eflicient construction than previously known However, one disadvantage of this type of construction does reside in the necessity for providing strength and therefore weight at certain points, thus aflecting efliciency. The reason for providing weight-adding, strengthening structure or hoop 6 at throat area 2 is because of the stresses that occur in this region tending to open up the throat area due to constriction of passing gases. Any structural deflection in this area is undesirable because the area of the throat must be maintained within critical limits. Even relatively small variations can result In addition, there may be high thermal stresses due to hot spots in the throat area resulting in differential deflections which are obviously undesirable. It may be added that while the throat area may be made initially smaller to compensate for pressure and thermal eflects, the design problems are diificult, at best. Thus, the additional structure or hoops are used to hold down any deflection.

Likewise, plug element 3 must be constructed so as to Withstand the stresses involved. In addition, cooling at the throat area 2 is a problem. If passages are provided for a coolant, strength of the section is diminished.

The disadvantages of increased weight as illustrated in FIG. 1 are obviated by the structure shown in FIG. 2 taken with FIG. 5 which form one embodiment of this invention. As shown in FIG. 2, the combustion chamber is defined by hoop segments 10. These segments are hollow with a converging portion 11a and a diverging portion 11b forming a narrower throat 11. This throat may be formed by pinching, crimping or by other conventional methods. Member 19 is an injector element. When all of the hoop segments are assembled, a toroidal chamber structure such as shown in FIG. 5 is formed. To form such a toroidal chamber, it is preferable to form each hoop with a slightly larger cross-sectional area at the outside portion of the chamber as at 10a which tapers to a smaller cross-sectional area at the inside portion of the chamber as at 10b. This allows full circumferential contact between adjacent hoops when they are assembled in toroidal position.

Preferably the hoops are formed of tantalum which is able to withstand high temperatures and remain ductile at cryogenic temperatures encountered during cold soak prior to engine firing. However, other materials dependent on the particular temperature and strength requirements are within the scope of this invention and include Permanickel which is a wrought, age hardenable nickel alloy containing approximately 98.5 percent of nickel, and molybdenum.

Referring now to FIG. 3, the operation of the device will be described. For simplicity, this description is directed to a rocket engine utilizing a bi-propellant system. That is, fuel and oxidizer is used although monopropellants can also be employed using the described invention. In FIG. 3, fuel such as hydrogen is supplied by means such as pumps or manifolds into inlet 13 of injector 19 with the major portion of the fuel passing in a circular fashion through hoop segment 10. At area 11, the fuel increases speed because of the narrow section 11. The fuel then passes into manifold area 14, through ports 15, and into combustion chamber 12. Port 4 is optionally provided to inject fuel directly and tangentially into the combustion chamber. This provides additional cooling of the hoops and aids in the mixing of propellants. Simultaneously, oxidizer such as liquid oxygen is introduced into manifold area 16 (see FIG. 4), ports 17 and likewise into combustion chamber 12. After combustion, the gases are passed through throat area 18 (see FIG. 2), expand, and are deflected by skirt 30 (FIG. resulting in thrust. Suitable propellant lines (not shown) are connected to inlet 13 and to manifold area 16. Although not shown, any conventional means may be employed to initiate combustion in the combustion chamber. Such means may include conventional pyrotechnic igniters, glowplugs, or spark plugs. Further, the reactants, i.e., fuel and oxidizer, may be such as to hypergolically or spontaneously react upon contact.

Two advantages result in this operation. Firstly, the combustion chamber is regeneratively cooled by passing the reactants through the combustion chamber tubes 10. This results in not only cooling the walls of the combustion chamber, but in heating reactants closer to combustion temperatures. The second advantage resides in the saving of weight. The tubes resist any tendency for separation of throat area 18; that is, tubes undergo longitudinal tension stresses in the region 11 and therefore resist separation. Thus, it can be seen that the necessity for heavy supporting structure at this point has been eliminated.

Another embodiment of this invention is shown in FIG. 6. As in the previous embodiment, a fuel such as liquid hydrogen or a conventional hydrocarbon J-P fuel is led into injector 20 by conventional means not shown, passes through tube or hoop segment 21, increases in velocity at section 22 due to the narrow passage area, decreases velocity in portion 24, increases velocity at portion 25 and is injected through injector 20 into combustion cham ber 26. As contrasted with FIG. 2, injector 20 is shown as integral rather than segmented. Liquid oxygen or other oxidizer is likewise fed from a conventional manifold or pump system into injector 20 and is combusted with the fuel in chamber 26. The combustion gases pass through throat areas 27 and 28 formed by narrowed portions 22 land 25 into expansion chamber 29. The gases passing through throat area 27 are deflected by skirt portion 301: while the gases passing through throat area 28 are deflected by a spike portion 31 of the nozzle. This results in axial thrust. The purpose of section 31 is to eliminate center plug drag. It is referred to as a spike because of its shape. It should be noted that in all embodiments, the narrow portion of each hoop segment not only provides for a throat area to allow passage of the combusted propellants, but also provides for increased cooling at the throat area. This is accomplished by the increase in the velocity of the propellant which results in a greater heat transfer from the throat area to the propellant. Although not shown, it is within the scope of this invention to provide additional combustion chambers. In such a modification,

concentric toroidal chambers would be used in conjunction with a plurality of combustion chambers 29 arranged concentrically.

Referring now to FIG. 7, a further modification of the invention is shown. Fuel such as liquid hydrogen is introduced into manifold 40 by pumps or other means, passes into tube members 41 and (as shown in FIG. 8) into injector 42. The fuel then passes into combustion chamber 43. Likewise, liquid oxygen, for example, passes from manifold 44, fed by pumps or other means not shown, into hoop segment or tube 45 into injector 42, then into combustion chamber 43 as in previous embodiments. Throat areas 46 and 47 are provided as in previous illustrations to allow the combusted gases to pass into expansion chamber 48 and chamber 49 (see FIG. 8). An important feature of this embodiment is the provision of hoop segments as the deflection portion of the nozzle rather than separate skirt portions. This not only allows the skirt portions to be cooled by the passage of propellants, but eliminates the need for additional skirt portions. FIG. 9 illustrates a view partially in cross section of FIG. 8 and discloses how the tubes 41 and 45 bypass each other.

FIG. 10 discloses a further embodiment of this invention. In this embodiment, an oxidizer manifold 50 supplied by conventional means (not shown) passes oxidizer into passage 51 of injector 52 into and through a core member 53 and then manifold area 54 of injector 52. The oxidizer then passes into combustion chamber 55 by means of ports 56. Similarly, fuel from manifold 57 passes into area 58 formed between core 53 and outer hoop segment 59. This embodiment has two main advantages; not only is there a heat transfer between the fuel and oxidizer due to the concentric tubes, but the cores 53 provide additional strength to the hoop segment combustion chamber. The fuel after passing through the annular volume formed by hoop segment 59 and core 58 passes into manifold area 60, passages 61 and into combustion chamber 55. The throat area is similar to previous embodiments and is formed by narrowed portion 62 of hoop segment 59 surrounding narrow portion 63 of core 53. The arrows shown represent the propellant flow.

A further modification is shown in FIG. 11. Instead of converging and diverging sections in the hoop segments, a relatively long narrowed portion is formed in each hoop segment. This is shown as 70 in hoop segments 71 and 72 in hoop segment 73. This forms throat portions 74. The advantage of this modification is that the throat area can be varied in area by merely changing the length of the narrowed portions or the degree of bend in adjacent hoop segments. In the modification illustrated, 72 is depicted as a straight portion whereas narrowed portions 70 are shown as a continuation of the curvilinear hoop segment 71. Thus it can be seen that by varying the degree of bend in adjacent outer portions 70 and 72, the area of the throat can be accordingly varied. FIG. 12 taken along the line of 1212 of FIG. 11 illustrates a narrow portion of each throat area. Likewise, FIG. 13 taken along the line of 1313 of FIG. 11 shows a wider throat area for the passage of combustion gases illustrated in both views by numeral 75. It is further within the scope of this invention to provide that selected hoops may have portions that loop outwardly so as to provide a throat area.

Since tube dimensions must be thick enough so as to resist tensile stresses while at the same time be maintained at a minimum dimension for eflicient heat transfer, a further modification is illustrated in FIG. 14. In this modification, tube or hoop segment is maintained at a thin dimension for eflicient heat transfer. In addition, internal ribs 81 are provided which function advantageously in two ways. Firstly, they reinforce hoop segment 80 and secondly, function as an additional area for efficient heat transfer. It is further within the scope of this invention to provide that said ribs extend outwardly rather than inwardly as shown. The ribs may be formed by extrusion or by forming them on a flat section and rolling the section into a tube.

Thus, it can be seen that in all of the embodiments, an improved injection and combustion assembly has been provided. The use of hoop segments result in not only efiicient cooling, but afford a savings in weight due to all stresses taken up in the longitudinal sections forming the hoop segments.

Although not shown, the throat areas formed between adjacent hoop segments can be varied in such a way as to vary the angle of thrust of the combusted gases as it leaves the combustion chamber. This may be accomplished by modifying the constrictions in the hoops so as to change the direction of combusted gases in a nonradial direction such as tangentially. The throat areas further can be formed by aerodynamic fins or the like on the hoop segments so as to vary the angle of the escaping combustion gases. This is particularly useful for directional control.

Although described as a bi-propellant system in all modifications, it is further apparent that this system can be used for mono-propellant systems or, for that matter, for any multi-propellant system. For example, in a mono-propellant system, the propellant would be passed around the hoop segments and injected at the injector without the need for additional propellants.

It is further within the scope of this invention to utilize a coolant other than fuel or oxidizer in this hoop. In all embodiments the propellants may be injected directly and a separate coolant forced through the hoop. This coolant could either be recirculated or disposed of as it is used.

In the embodiment illustrated in FIG. 10, this auxiliary coolant could be forced through the space between the hoop and core with a propellant in the core.

Although the throat areas in all embodiments are shown as being in the same location from one segment to the next, it is within the scope of this invention to vary the number, spacing and location. For example, various throat areas may be staggered or eliminated dependent on the desired angle of thrust, the degree of thrust or to eliminate problems of turbulence. The spacing location and size of the throats may also be adapted to accommodate different nozzle designs.

Although disclosed as applied to a rocket engine, it is within the scope of this invention to use the structure and concept in other applications such as a standing propulsion unit or power plant. However, these are examples only and are not intended to be all inclusive.

Also within the scope of this invention is the use of one hoop shaped conduit in conjunction with other means such as plates on either side of the hoop to form a combustion chamber. Also, instead of a toroidal chamber, it is contemplated that a cylindrical combustion chamber could be utilized comprising a plurality of hoops forming the chamber.

It therefore becomes evident that the instant invention solves a problem in the prior art in an expeditious manner. Not only is the combustion chamber regeneratively cooled by the propellant, but the propellant is additionally heated to provide eflicient combustion. In addition, the narrow throat area in all embodiments provides for an increased velocity of the coolant, thus providing a more efficient heat transfer in the throat region whereas hot spots occur due to the escape of combustion gases. Also, a savings in weight is accomplished because of the segmented hoop construction.

Although the invention has been described in detail, it is to be clearly understood that the same is by way of illustration and example only, and is not to be taken by way of limitation, the spirit and scope of this invention being limited only by the terms of the appended claims.

We claim:

1. A modular rocket engine including a combustion chamber comprising a series of abutting, juxtaposed, substantially duplicate segments; each segment comprising an injector portion,

coolant tube loops extending from opposite sides of said injector portion, said loops terminating in a throat-forming portion spaced from and facing said injector portion,

manifold means connected to one of said loops and means to flow coolant through said tube loops from said manifold means to said injector portion, each said segment when in abutting position with other segments being adapted to conduct coolant therethrough in parallel fiow paths, the inwardly facing portions of said loops and the injector portion in all of said segments collectively forming and defining said combustion chamher.

2. The invention as set forth in claim 1, in which said throat-forming portions comprise sections of said loops having reduced cross-sectional area such that a throat means is formed between the loops of abutting segments.

3. The invention as set forth in claim 1, in which said segments extend peripherally around an annulus and wherein the loops vary in cross-sectional area so as to allow full circumferential contact between abutting loops in the juxtaposed segments.

4. The invention as set forth in claim 1 wherein a plurality of throat-forming portions are formed by said loops to allow combustion gases to exit in a plurality of directions.

5. The invention as set forth in claim 1 wherein a nozzle is provided to deflect combustion gases exiting from said combustion chamber through said throat-forming portions.

6. A liquid fuel rocket engine comprising a combustion chamber, said chamber being formed of a plurality of juxtaposed curvilinear hollow conduits defining a substantially closed combustion chamber volume,

injector means in communication with said conduits and oriented to inject propellant into the chamber volume defined by said conduits, said combustion chamber having a longitudinal center axis at right angles to the direction of propellant injection, aperture means bounded by portions of said hollow conduits opposite said injector means forming a throat section for exiting of gases resultant from combustion of injected propellant and ejection of said gases from said throat section to produce useful thrust, said combustion chamber longitudinal center axis being at right angles to the flow vector of said gases exiting from said throat.

7. A rocket engine comprising;

an injector,

a propellant source,

an elongated combustion chamber having a longitudinal axis of symmetry, a throat section, and

an exhaust nozzle communicating with said combustion chamber, said injector being adapted to inject propellant from said source into said combustion chamber to produce combustion gases, said longitudinal axis being normal to the direction of flow of combustion gases through said throat section, said combustion chamber being substantially defined by a plurality of curved conduits connected in flow parallel relationship between said propellant source and said combustion chamber, each of said conduits lying in a plane normal to the longitudinal axis of said combustion chamber, said throat section being formed between juxtaposed portions of said conduits and being cooled by flow of propellant through said conduits.

(References on following page) 7 8 References Cited by the Examiner 3,127,737 4/ 1964 Ledwith 60-355 UNITED STATES PATENTS 3,127,738 4/1964 Hasbrouck et a1. 60-35.6

4 5/1959 Coty 60 35.6 3,13 ,224 5/196 Llpplncott et a1 60 356 6/1963 Panlla X 5 MARK NEWMAN, Primary Examiner. 8/1963 Newcomb 6035.6 12 19 3 Adamson 0 35 5 UEL FEINBERG, G. L. PETERSON, Examiners.

Claims (1)

1. A MODULAR ROCKET ENGINE INCLUDING A COMBUSTION CHAMBER COMPRISING A SERIES OF ABUTTING, JUXTAPOSED, SUBSTANTIALLY DUPLICATE SEGMENTS; EACH OPPOSITE SIDES OF AN INJECTOR PORTION, COOLANT TUBE LOOPS EXTENDING FROM OPPOSITE SIDES OF SAID INJECTOR PORTION, SAID LOOPS TERMINATING IN A THROAT-FORMING PORTION SPACED FROM AND FACING SAID INJECTOR PORTION, MANIFOLD MEANS CONNECTED TO ONE OF SAID LOOPS AND MEANS TO FLOW COOLANT THROUGH SAID TUBE LOOPS FROM SAID MANIFOLD MEANS TO SAID INJECTOR PORTION, EACH SAID SEGMENT WHEN IN ABUTTING POSITION WITH OTHER SEGMENTS BEING ADAPTED TO CONDUCT COOLANT THERETHROUGH IN PARALLEL FLOW PATHS, THE INWARDLY FACING PORTIONS OF SAID LOOPS AND THE INJECTOR PORTION IN ALL OF SAID SEGMENTS COLLECTIVELY FORMING AND DEFINING SAID COMBUSTION CHAMBER.

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