US3451221A - Supersonic combustion nozzle - Google Patents

Description

June24, 1969 R. o. BOE

- SUPERSONIC COMBUSTION NOZZLE Sheet Filed July '26, 1966 /PO, A/A/ 0. 505,

INVEN'IOR.

United States Patent M 3,451,221 I SUPERSONIC COMBUSTION NOZZLE Rollin 0. Boe, Canoga Park, Califi, assignor to The Marquardt Corporation, Van Nuys, Calif., a corporation of California Filed July 26, 1966, Ser. No. 568,029 Int. Cl. F02g 3/00; F021: 9/00, 1/00 US. Cl. 60-258 14 Claims ABSTRACT OF THE DISCLOSURE A convergent-divergent supersonic combustion nozzle in which high pressure oxidizer is introduced subsonically to the convergent portion for expansion to supersonic velocity and fuel is introduced in the vicinity of the throat parallel to the oxidizer flow for supersonic combustion therewith in the divergent portion. The fuel and oxidizer may be hypergolic, or combustion may be initiated by electric or pilot burner means.

This invention relates to a supersonic combustion nozzle and more particularly to a nozzle having burning in the supersonic portion of the nozzle thereby eliminating heating problems in the region of the sonic throat.

In present conventional rockets, combustion takes place in a subsonic chamber and the resulting hot gas flow is exhausted through a sonic throat and a supersonic nozzle. The heat transfer to the nozzle is maximum in the region of the sonic throat since the heat transfer is largest where the diameter is minimum. The high temperatures encountered with high Mach number nozzles have resulted in a serious heating problem of the material used to make the sonic throat. In order to keep the throat and the subsonic combustion chamber cool, water jackets and cool air boundary layers must be utilized. These methods of cooling take heat from the main gas flow, resulting in reduced efficiency.

The supersonic combustion nozzle of the present invention utilizes burning in the supersonic stream of the nozzle in order to reduce the cooling requirement in the critical sonic portion of the nozzle and to reduce the size of the sonic throat. Also, the need for large subsonic combustion chambers is eliminated and most of the heat of combustion remains in the supersonic exhaust flow. The nozzles have application both as propulsion nozzles and as flight simulation test nozzles. An electrical arc can be utilized to accomplish ignition of the fuel and oxidizer in the supersonic region or a small pilot burner located in the supersonic region can be utilized. Also, the fuel can be introduced directly into the supersonic oxidizer flow stream for auto ignition when compatible fuel and oxidizers are utilized. In most cases, the fuel injection system will introduce fuel in the vicinity of the sonic throat and the fuel is added to the oxidizer in the sonic region either through nozzles in the flow stream line or through the walls of the nozzle. The nozzle configuration for supersonic combustion differs from the standard nozzle configuration since large expansion will take place during supersonic combustion requiring larger expansion ratios than encountered in normal supersonic nozzles.

It is therefore an object of the invention to provide a supersonic combustion nozzle in which combustion occurs in the supersonic portion of the nozzle thereby eliminating heating problems in the region of the sonic throat.

Another object of the present invention is to provide a supersonic combustion nozzle in which the fuel and oxidizer are mixed together and combusted in the supersonic flow portion of the nozzle, thereby reducing the cooling requirements in the sonic portion of the nozzle 3,451,221 Patented June 24, 1969 and eliminating the need for large subsonic combustion chambers.

Another object of the invention is to provide a supersonic combustion nozzle in which fuel and oxidizer are mixed in the vicinity of the supersonic portion of the nozzle and are combusted by ignition means located in this region.

These and other objects of the invention not specifically set forth above will become readily apparent from the accompanying description and drawings in which:

FIGURE 1 is a section through a supersonic combustion nozzle of the present invention showing the air and fuel manifolds;

FIGURE 2 is an enlarged transverse section along line 22 of FIGURE 1 showing the fuel tubes within the nozzle throat;

FIGURE 3 is a section of a modified supersonic combustion nozzle in which the oxidizer is introduced to indi- Vidual tubes, each containing a fuel supply tube;

FIGURE 4 is a transverse section along line 4-4 of FIGURE 3 showing the oxidizer manifold;

FIGURE 5 is a section of a third modification of the supersonic combustion nozzle showing the fuel tubes inserted through the sides of the nozzle and the electrodes located in the supersonic flow region;

FIGURE 6 is a transverse section along line 66 of FIGURE 5;

FIGURE 7 is a sectional view of a fourth modification of the supersonic combustion nozzle in which the fuel and oxidizer flow through separate nozzles and auto-ignite in the supersonic region;

FIGURE 8 is a transverse section along line 8-8 of FIGURE 7 showing the manifolding for the fuel; and

FIGURE 9 is a schematic of the nozzle showing the various regions of the nozzle.

Referring to FIGURES 1 and 2, the supersonic combustion nozzle comprises a nozzle block 10 having an entrance '11 leading to the nozzle throat 12. The diverging portion 13 of the nozzle is connected at end 13a to the nozzle block 10 and the other end 13b carries a flange 14. A cover 15 extends from the nozzle block 10 parallel to the diverging portion 13 and has a step section 16. Spiral spacer rings 17 are connected to the diverging nozzle portion 13 to form an annular cooling passage 17a be tween the diverging nozzle portion 13 and the cover 15. A cylinder -18 projects from the flange 14 and has an end 18:: received in an annular cup 19 containing a seal 20 in the form of a flexible gasket. The cup 19 is secured to a larger annular member 21 and a plurality of bolts 22 project between the member 21 and a circular flange 23 carried by the cylinder 18. Tightening of the bolts 22 will cause the end 1811 to squeeze the gasket 20 into sealing relationship with the step section 16 while still permitting relative movement of the cylinder 18 and the section 16. Since the spacer rings 17 are secured only to the diverging nozzle portion 13 and since the gasket 20 can move relative to the offset cover section 16, it is apparent that a temperature differential will not produce stresses in this nozzle structure.

A casing member 30 has a flange 31 which is connected to a flange 32 on the nozzle block 10 by means of a plurality of bolts 33. The flange 31 is located at the end of a cylindrical portion 35 of the casing member and portion 35 connects with a conical portion 36 leading to a second cylindrical portion 37. A pilot combustion chamber 40 is located within the casing member 30 and is spaced therefrom by means of spacer rings 41 connected only to the chamber 40. A base plate 43 closes the end of casing member 30 and also supports and closes the end of chamber 40. The base plate 43 supports an igniter 44 of any well known standard construction and fuel passage 45 and oxidizer passage 46 discharge adjacent the igniter in order to produce combustion within the pilot combustion chamber. The reduced discharge end 47 of combustion chamber 40 passes through the nozzle 12 and opens into the region of the nozzle where the flow is supersonic in order to provide a small pilot burner. A small amount of fuel and oxidizer burn subsonically in a combustion chamber 40 and the combustion products flow through the pilot burner 47 at a high enough temperature to ignite the cold oxidizer and fuel flows to the main nozzle.

A circular manifold 50 surrounds the cylindrical portion 35 of easing member 30 and connects to an entrance passage 51 which supplies high pressure oxidizer, such as air or oxygen, to the manifold. The manifold has four radial passages 52 (only two of which are shown) which lead to a space 53 located between the nozzle block and a partition 54 in casing portion 35. The high pressure oxidizer supplied to space 53 flows through the nozzle throat 12 and obtains a supersonic velocity in the region of 55 downstream of the throat. A smaller circular fuel manifold 56 also surrounds the cylindrical casing portion 35 and twelve radial passages 57 lead from the manifold through the casing portion 35 and project through the nozzle throat 12 so that the end 57a of each passage is located adjacent the end of pilot burner 47. Fuel is supplied to manifold 56 through the passage 60 leading to a fuel supply source.

In operation of the nozzle, oxidizer at high pressure flows from manifold 51 through the space 53 and through nozzle throat 12 and has supersonic velocity in the nozzle region 55. However, the temperature and static pressure of the oxidizer have been considerably reduced at region 55. For instance, air introduced at about 400 psi. into the manifold 50 would be at about 50 p.s.i. in the region 55 and at a temperature of -170 F. However, the air has reached a velocity of about 1700 feet per second corresponding to Mach 2. The static pressure of the fuel which is introduced into the supersonic region 55 through the fuel tubes 57 will be substantially the same as the static pressure of the oxidizer in region 55. In order to prevent shock waves from building up when the hot flow from the pilot combustion chamber 40 meets the supersonic cold flow, the velocity and static pressure of the hot flow from chamber 40 is made substantially the same as the velocity and static pressure of the supersonic cold flow. This is accomplished by having the hot flow enter the region 55 at a subsonic velocity (about Mach 0.5) and at a temperature (about 5000 F.) which will provide the same static pressure and velocity as the cold flow.

Since the pilot combustion chamber operates at a low combustion pressure, the thick walled chambers are not required to contain the pressure. Also, the ratio between the oxidizer flow and the hot flow from the combustion chamber 40 can be in the neighborhood of 50 to 1. Therefore, high volumes of flow from the combustion chamber 40 are not required. Since the pilot burner tube 47 of the pilot combustion chamber is parallel to the supersonic cold flow through the nozzle, reduction of shock waves in the supersonic flow is obtained and more uniform combustion results from starting the burning at the center of the flow, A schematic illustration of the various regions in the nozzle is illustrated in FIGURE 9. Subsonic flow exists in region 1 upstream of the throat 12 and the first expansion of the oxidizer to a predetermined Mach number for burning will occur in the region 2. In region 3, constant Mach number combustion will occur and in region 4, the final expansion to the maximum Mach number will result.

A modification of the supersonic combustion nozzle is illustrated in FIGURES 3 and 4. The nozzle has throat portion 65 connected with a diverging section 66. The entrance end 67 to the nozzle is connected by means of a scalloped band 64 to the ends 68a of a plurality of circular oxidizer tubes 68 which are welded together at their abutment location 69 to form a circular configuration. The entrance ends 68b of the tubes 68 are connected to an annular space 70 defined by sides 72 and 73 and an annular base 74. An annular oxidizer manifold 75 surrounds the tubes 68 and is connected by passage 76 to a high pressure oxidizer supply. Four radial passages 78 connect manifold 75 to space 70. Thus, oxidizer from manifold 75 is conducted through the twelve tubes 68 to the entrance of the nozzle. The oxidizer then passes through the nozzle throat 65 and is supersonic in the region 80.

The interior surface of the tubes 68 form the interior surface of pilot combustion chamber 81 which produces a hot gas flow at the end 83 for igniting the main supply of fuel and oxidizer. The pilot burner end 83 of the combustion chamber 81 is attached to the interior of the ends 68a of tubes 68 by a scalloped band 83a (not shown) and projects through the throat 65 of the nozzle into the supersonic cold flow. An igniter 84 of standard construction is located in the back wall 85 which closes the end of chamber 81 and fuel line 86 and oxidant line 87 terminate in the vicinity of the igniter in order to produce hot gas by combustion. The small flow of hot gas at the exit end 83 is of sufficiently high temperature to ignite the main oxidant and fuel supply to the nozzle. A fuel tube 91 extends through the center of each of the oxidizer tubes 68 and the ends 91a of the tubes 91 are located adjacent the end 83 of the pilot burner. The other ends 91b of the fuel tubes extend through side 71 and connect with an annular fuel manifold 95 which is connected to a source of fuel supply through the passage 96.

The operation of the modification of FIGURES 3 and 4 is similar to that of the first embodiment since high pressure oxidizer is introduced to each of the tubes 68 and passes through the nozzle throat 65 to become supersonic in the region 80. At this location, hot gas flow is introduced from the pilot combustion chamber 81, and also fuel is introduced through the individual tubes 91. The resulting mixture of fuel and oxidizer is ignited by the hot gas fiow in the supersonic region 80. In this modification, the individual air passages 68 are connected together to provide an annular combustion chamber 81 without additional structure so that a light nozzle motor is provided. The oxidizer flow through tubes 68 cools the portion of the tubes 68 which form the interior surface of the pilot combustion chamber 81. Also, the exterior surface of the tubes 68 provide the outer surface of the nozzle and additional cowlings and structure are not required Referring to a third modification shown in FIGURES 5 and 6, the nozzle has an entrance portion and a diverging portion 101 connected together by the throat portion 102. A plurality of fuel tubes 103 extend through the entrance end wall of the nozzle and their ends 103a terminate in the vicinity of the throat of the nozzle. The other end 1031; of each tube connects with a suitable fuel manifold (not shown) and an oxidizer is introduced under high pressure into the entrance portion 100 of the nozzle from an oxidizer source. A pair of electrodes 104 and 105 extend through the throat of the nozzle to produce an electric are 106 within the supersonic region 107 of the nozzle. The electric arc is utilized to ignite mixtures of fuel and oxidizer which are not self-igniting. Since the electrodes are inserted parallel to the flow, they produce minimum effect on the supersonic air stream and the supersonic flow will serve to cool the electrodes. Two alternate locations for the ends 103a are designated by the positions A and B in FIGURE 5. In the case of fuels and oxidizers which can be mixed without auto-ignition, the ends of the tubes can be in the subsonic region ahead of the sonic throat as indicated by position A of the nozzle ends. However, for many fuel-oxidizer combinations, this is not possible and therefore the nozzle ends must be located at least as far downstream as position B to introduce the fuel in the vicinity of the sonic throat. When the fuel nozzles are located in position A upstream of sonic throat, they allow the maximum time for the fuel and oxidizer to mix.

In the case of fuel-oxidizer combinations that autoignite, the etfective area of the sonic throat can be reduced to the extent that combustion takes place ahead of the sonic throat by placing the nozzle exit ends upstream of the throat. Thus, the effective area of the nozzle can be changed by moving the fuel nozzle to vary the amount of combustion which takes place before reaching the throat. The change in efiective nozzle area can result in a change of the effective area ratio between the sonic throat and the supersonic exit nozzle, thereby serving as a means for varying Mach number of a fixed configura tion nozzle. The Mach number can also be varied by changing the fuel-oxidizer ratio since the amount of expansion due to supersonic combustion will vary. Generally, the Mach number goes down as the fuel oxidizer ratio increases. Thus, the third modification of the invention provides a convenient structure in which to vary the location of the outlet ends of the fuel tubes in order to provide a large variety of conditions for the fueloxidizer combustion and this embodiment can also employ electric arc type of igniter.

The fourth embodiment of the invention is illustrated in FIGURES 7 and 8 wherein the fuel and oxidizer are introduced through individual annular nozzles into supersonic layers of mixture which will auto-ignite. The nozzle consists of a divergent portion 115 which is connected by throat portion 114 with an entrance portion 116 through which the oxidizer flows. Four annular sleeves 117 are located in the entrance portion and each sleeve terminates in an enlarged end 117a to form a nozzle 119 with its adjacent sleeve end 117a and a nozzle 119a with throat 114. The individual sleeves are attached to four passages 130 leading from a fuel manifold passage 131 which extends transversely across the entrance portion 116. The outer annular passage 120 receives oxidizer from entrance 116 and discharges the oxidizer through nozzle 119a. The passages 121 and 122 also receive oxidizer and each passage discharges through a nozzle 119. The annular passages 125 and 126 each receive fuel from two passages 130 and passages 125 and 126 each dis charge through a nozzle 119. The passages 125 and 126 are blocked from the oxidizer in entrance portion 116 by the back plates 133 and 134 which contain openings for passages 130. In the fourth modification, both the fuel and oxidizer are discharged in annular layers and at supersonic velocities into the divergent section of the nozzle. Fuels and oxidizers are utilized that will autoignite upon contact without additional ignition means and the alternate positioning of the layers of fuel and oxygen will facilitate the auto-ignition. It is understood that the number of fuel and oxidizer nozzles can be varied in order to obtain the proper flow of these substances to provide the desired fuel-oxidizer ratio.

By the present invention there is provided a supersonic combustion nozzle in which high pressure, cold flow passes through a sonic throat to reach supersonic velocity and the cold flow is thereafter mixed with the fuel so that combustion can take place in the supersonic region of the nozzle. Ignition can be initiated in the supersonic region either electrically or by means of a hot gas from a pilot burner. While it is usually desirable to introduce the fuel into the supersonic region of cold flow, the point of introduction of the fuel can be varied as indicated to produce various modifications of nozzle performance, such as permitting a small amount of the combustion to take place in the subsonic region and thereby vary the effective area of the sonic throat. Since substantially all combustion takes place in the supersonic region, only a small amount of cooling is required in the throat portion of the nozzle and no large subsonic combustion chamber is required. In general, supersonic combustion will require larger expansion ratios than encountered in present subsonic nozzles. Also, the length of the supersonic combustion chamber can be minimized by burning at low supersonic Mach numbers.

Various other modifications in addition to those described herein are contemplated by those skilled in the art without departing from the spirit and scope of the invention hereinafter defined by the appended claims.

What is claimed is:

1. A supersonic combustion nozzle comprising;

a nozzle throat portion connecting with a convergent nozzle entrance portion and discharging through a divergent nozzle portion;

means for introducing high pressure oxidizer subsonically to said entrance portion for expansion through said throat portion to reach supersonic velocity in said divergent portion; and

means for introducing fuel into said nozzle substantially parallel to the oxidizer flow in the vicinity of said throat portion to produce supersonic combustion in the divergent portion of the nozzle.

2. A supersonic combustion nozzle as defined in claim 1 wherein said fuel introducing means comprises a plurality of fuel tubes having their outlet in the supersonic flow region of the nozzle.

3. A supersonic combustion nozzle as defined in claim 2 having ignition means located in the vicinity of the outlets of said fuel tubes in said supersonic region for initiating combustion downstream of the nozzle throat portion.

4. A supersonic combustion nozzle as defined in claim 3 wherein said ignition means comprises a subsonic combustion pilot burner to produce a hot flow for igniting said fuel and oxidizer.

5. A supersonic combustion nozzle as defined in claim 4 wherein said pilot burner comprises a tube axially positioned at the center of said throat portion and extending upstream of said throat to connect with said pilot combustion chamber.

6. A supersonic combustion nozzle as defined in claim 3 wherein said ignition means comprises electrodes eX- tending axially of said throat portion and terminating in the vicinity of the supersonic oxidizer flow.

7. A supersonic combustion nozzle as defined in claim 1 wherein said fuel introducing means comprises a plurality of fuel tubes having their outlets slightly upstream of said throat portion to allow maximum time for fuel and oxidizer to mix, said fuel and oxidizer being of the type which do not auto-ignite.

8. A supersonic combustion nozzle as defined in claim 1 wherein said fuel introducing means introduces said fuel slightly upstream of said throat to vary the elfective area of the nozzle; said fuel and oxidizer being of the type which auto-ignite.

9. A supersonic nozzle as defined in claim 4 wherein said oxidizer introducing means comprises a plurality of circular tubes attached together in a circular configuration and discharging into said throat portion of said nozzle, the interior surfaces of said tubes forming said pilot combustion chamber, said fuel introducing means comprising individual fuel tubes extending through each of said oxidizer tubes.

10. A supersonic combustion nozzle .as defined in claim 4 wherein said throat portion comprises a plurality of individual nozzles discharging into said divergent portion, said oxidizer and fuel introducing means comprising a plurality of passages connecting some of said nozzles to oxidizer and other of said nozzles to fuel.

11. A supersonic combustion nozzle as defined in claim 2 wherein said fuel tubes extend through said throat portion.

12. A supersonic combustion nozzle as defined in claim 4 wherein said oxidizer introducing means introduces said oxidizer at a pressure to produce supersonic flow downstream of said throat portion at a reduced static pressure, said pilot burner producing a hot flame at a velocity and static pressure matching that of supersonic oxidizer in order to prevent the development of shock waves in the region of mixing of the fuel and oxidant.

13. A supersonic combustion nozzle as defined in claim 2 wherein said fuel introducing means introduces said fuel to said oxidizer at the static pressure of said oxidizer.

14. A supersonic combustion nozzle as defined in claim 1 having ignition means located in the region of supersonic velocity.

References Cited UNITED STATES PATENTS 2,992,527 7/1961 Masnik 60-270 3,095,694 7/1963 Walter 60261 3,112,988 12/1963 Coldren 60-270 9/1964 Rocca 60261 10/1966 Sippel 60261 10/1966 Dugger 60-270 5/1967 Walter 60-261 6/ 1967 Sanger 60270 8/1967 Rhodes 60270 FOREIGN PATENTS 1/1966 Germany.

US. Cl. X.R.

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