US2774216A

US2774216A - Rocket motors

Info

Publication number: US2774216A
Application number: US319832A
Authority: US
Inventors: Allen Sidney
Original assignee: Armstrong Siddeley Motors Ltd
Current assignee: Armstrong Siddeley Motors Ltd
Priority date: 1952-01-21
Filing date: 1952-11-12
Publication date: 1956-12-18
Anticipated expiration: 1973-12-18
Also published as: BE552192A

Description

Dec. 18, 1956 Filed Nov. 12, .1952

s'. ALLEN 2,774,216

ROCKET MOTORS 3 Sheets-Sheet 1 Dec. 18, 1956 Filed Nov. 12, 1952 S. ALLEN ROCKET MOTORS 3 SheetS -Sheet 2 Dec. 18, 1956 Filed NOV. 12, 1952 s. ALLEN ROCKET MOTORS 3 Sheet-Sheet 3 United States Patent ROCKET MOTORS Application November 12, 1952, Serial No. 319,832

Claims priority, application Great Britain January 21, 1952 3 Claims. (Cl. 6035.6)

This invention relates to a rocket motor of the kind having a combustion chamber designed to operate on a liquid fuel and a liquid oxidizer as propellent.

When the liquid propellent combination is one giving a very high flame temperature, particularly a combination using liquid oxygen as the oxidizer, it is important that the volume of the combustion zone of the combustion chamber should be kept as small as possible, in order to reduce the total heat transfer to the walls of the combustion chamber, and thus to reduce the amount of the coolant necessary.

A means of achieving this, according to the invention, is to vaporize and heat the oxidizer upstream of the com bustion zone. In consequence the actual combustion process is accelerated, due to the oxidizer being preheated (preferably to a temperature which gives spontaneous ignition with the fuel)i. e., the time of the combustion process is less than if the oxidizer were not preheated. In other words, the combustion process can be completed with a very short flame length for a given thrust, thus allowing the surface of the combustion chamber requiring cooling to be reduced.

The invention therefore consists in a rocket motor having a combustion chamber designed to operate on liquid oxidizer and a liquid fuel, and is characterized in that the whole or at least the main portion of the oxidizer is admitted to a preheating portion of the combustion chamber, upstream of the fuel supply thereto, and is subjected to a heating means by which it is evaporated and heated to a relatively high temperature before the fuel is added, which high temperature may be such as to cause spontaneous ignition with the fuel.

In the accompanying drawings:

Figure 1 is a plan of a rocket motor according to the invention;

Figure 2 is a sectional plan, to a larger scale, of the evaporator or preheating chamber;

Figure 3 is a sectional plan mainly of the upstream end of the combustion zone of the combustion chamber and showing one form of main fuel burner, the section of the latter being taken on the line 33 of Figure 4; and

Figure 4 is an elevation of the main fuel burner taken from the downstream end thereof.

It will be observed from Figure 1 that the preheating chamber, combustion chamber and nozzle of the motor are built up of a number of external rings having flanges which are secured together by bolts 11 or studs 12 namely, three

rings

14, 15 and 16 constituting the preheating chamber (see Figure 2), a transfer ring 17 at the downstream end thereof, and

rings

18 and 19 forming the combustion chamber, the latter terminating with a divergent nozzle the outline of which is shown at 20. A main fuel burner 21 is disposed in the upstream ring 18 of the main combustion zone.

The ring 19 of the main combustion zone has on its inner face a helical ridge 22, the apices of the helical ridge being engaged by a heat-resistant liner 23 for the actual combustion zone. A coolant is circulated in the jacket 2,774,216 Patented Dec. 18, 1956 space 24 round the liner 23, being, for example, taken in at the downstream end of the jacket space, as indicated at 25, and travelling upstream in a helical path determined by the helical ridge 22.

Incidentally, the ring 18 (Figure 3) is shown as having a liner 26 providing, with the ring 18, an annular space 27 round the upstream end of the liner 23. A coolant may be supplied to this space 27 under pressure to be passed through a number of small radial holes 28 in the liner 26 and form a film of coolant alongside the interior of the liner 23 to provide additional cooling for the latter;

No separate means for supplying the coolant to the annular space 27 is shown; but obviously leakage passages may, if desired, be provided from the outlet end 29 of the jacket space 24 past an internal flange 30, the associated gland 31 being omitted in this case, to allow of heated coolant from the jacket space 24 passing into the annular space 27. Any excess of coolant reaching the outlet end 29 may, in that case, be carried off through the outlet 32.

If water be the coolant, such excess may be discarded or returned for cooling and re-use; but. naturally, if for example liquid fuel be the coolant, the outlet 32 can be connected to the fuel supply pipe for the main burner which is hereinafter referred to. Naturally, too, different coolants may be used for the jacket space 24 and for the annular space 27 if desired.

In the construction shown the rocket motor includes, at the end remote from its outlet 20, the preheating portion into which the main supply of the liquid oxidizer is admitted, some of which is admitted as by means of supply pipes 34, 35 and annular passages 36, 37. Both of these annular passages communicate with

radial ducts

38 and 39, respectively, the former spraying the liquid oxidizer through a plurality of small, radial holes 41, and the latter spraying the oxidizer upstream through a plurality of small, axially-extending holes 42.

These supplies of the liquid oxidizer are admitted downstream of a heating means disposed adjacent the closed, upstream end 44 of the preheating chamber, for evaporating the liquid oxidizer and heating the resulting gases to a relatively high temperature before they are intimately mixed, preparatory to combustion, with the main supply of liquid fuel. In the present instance, the heating means is shown as including an injector or spray device 46 for a small portion of the liquid fuel, and an igniter plug 47, or use may be made of two

igniter plugs

47, 47a as shown. The remainder of the liquid oxidizer is delivered, by means not shown, to an annular chamber 49 whence it is sprayed downstream through holes 50 to impinge on a cone-shaped bafile plate 51. This may be provided with inclined holes 53 directing some of the liquid oxidizer towards the fuel sprayed by the spray device 46 to support the burning of this fuel The main portion of the oxidizer passing through the holes 50 passes round the edge of the battle 51, as indicated by the arrows 54, to cool the wall of the ring 15 whilst evaporating and becoming heated by the heated mixture in the middle of the ring.

The heated mixture then passes at high velocity through a constriction 55 where it picks up the further supply of the liquid oxidizer from the spray holes 41 and heats and evaporates it, the turbulence downstream of the constriction causing an intimate mixing of the heated gases with this further supply, and with the last supply of oxidizer from the spray holes 42, which latter supply is in consequence evaporated and heated. The mixture then passes through another constriction in the form of an annular bafile 56, which also promotes turbulence downstream in the heated mixture. Very thorough evaporation and uniformity of heating of the oxidizer therefore takes place upstream of the main burner 21, and, as

stated, the temperature is preferably such that spontaneous ignition of the fuel from the main burner will take place.

The main burner comprises inner and outer concentric rings 58, 59xwith annular passages 60, 60 in them terminating-downstream with fuel spray holes 61.- Disposed in the paths of these fuel spray holes 61 are

annular baffles

63, 64 which may be secured to the two burner rings as by-means ofscrews 65, 65. Fuelis led in by a piper67 to an annular space 68 round a liner portion 69 of the main burner and then distributed to the burner rings-58,- 59 by a seriesof angularly-spaced, radial, hollow bosses 70, 71, three of these being shown in Figure 4.

' Thus, the main supply of fuel-is distributed very uniformly over the cross-sectional area of the combustion chamber.

V What I claim as my invention and desire to secure by Letters=Patent of the United States is:

v '1. A rocket motor comprising a number of coaxial rings secured to one another and providing a preheating ehamber'upstream of the combustion chamber of the motor, the motor being designed to operate on a liquid fuel and a liquid oxidizer as propellent, a heating means at the upstream end of the preheating chamber including cans for supplying a small portion of the fuel and the oxidizer and an igniter to ignite-said small portions of fueland oxidizer, a constriction provided by the walls of the preheating chamber downstream of'the heating means, means for supplying a first portion of the rest of the oxidizer to spray holes in the wall providing the constriction, a further constriction in the preheating chamber downstreamof said first-mentioned constriction provided by a bafile, means for supplying the rest of the oxidizer to spray holes in said bafiie, a burner at the upstream end of the combustion chamber, and means for supplying the rest of the fuel to said burner.

12. A rocket motor having a combustion chamber designed to operate with supplies'of a liquid fuel and a liquid oxidizer, the combustion chamber providing a preheating and vaporizing chamber in its upstream end, means for admitting a main supply of the liquid fuel to the combustion chamber downstream of the preheating chamber, means for admitting at least the main portion of the liquid oxidizer to the preheating chamber, said main portion of the liquid oxidizer introduced into said preheating chamber at separate positions which are axially-spaced progressively in the downstream direction, a

temperature before'it encounters the main supply of fuel delivered to said combustion chamber whereby to accelerate the combustion of said main supply of fuel.

3. A rocket motor having a combustion chamber designed to operate with supplies of a liquid fuel and a liquid oxidizer, the combustion chamber providing a preheating and vaporizing chamber in its upstream end, means for admitting a main supply of the liquid fuel to the combustion chamber downstream of the preheating chamber, means for admitting at least the main portion of the liquid oxidizer to the preheating chamber, said main portion of the'liquid oxidizer introduced into said preheating chamber at separate positions which are axially-spaced progressively in the downstream direction,

a brittle at the upstream end of the preheating chamber serving to direct some of the liquid oxidizer along the wallof the initial part of the preheating chamber for cooling the'lattenmeans for admitting a small supply of the liquid fuel to the upstream end of said preheating chamber, an injection device for said small supply of liquid fuel, said baffle having fine holes through it for directing some of the oxidizer towards the injection device to support the combustion of said small supply of liquid fuehsaid small-supply of liquid fuel burning with the oxidizer supplied so as to vaporize the remainder of said portion of'said oxidizer andheat it to a relatively high' temperature before it encounters the main supply of fuel'delivered to said combustion chamber whereby to accelerate the combustion ofsaid main supply of fuel.

References Cited in the file ofthis patent UNITED STATES PATENTS 2,408,111 Truax et al. Sept. 24, 1946 2,523,656 Goddard Sept. 26, 1950 FOREIGN PATENTS 248,547 Germany June 25, 1912