GB1605332A - Improvements in Rockets - Google Patents

Description

(54) IMPROVEMENTS IN ROCKETS We, ROLLS-ROYCE LIMITED a British Company of 65 Buckingham Gate, London SW I E 6AT, formerly ROLLS ROYCE (1971) LIMITED of Norfolk House, St. Jame's Square.

London, SW I Y 4JS do hereby declare the invention, for which we pray that a patent may be granted to us, and the method by which it is to be performed, to be particularly described in and by the following statement: This invention relates to improvements in rockets and has particular reference to an integrated rocket ramjet.

The evolution of rocket designs and in particular missile designs has resulted in a trend towards dividing the rocket into two portions. a forward guidance and payload portion and a rearwardly disposed propulsion assembly.

A typical propulsion assembly comprises a rocket fuel propellant chamber and a ramjet combustion system disposed upstream of the chamber and operative, after burning of the rocket fuel propellant. with air received from an inlet or inlets.

The inlet or inlets may be stowed within the generally cylindrical body of the rocket to facilitate launching the rocket and to prevent the creation of excessive drag during acceleration of the rocket by the rocket fuel. In one rocket type four inlets are provided and, on reaching a sufficient speed, covers are blown off the intakes by small explosive charges and the inlets swing outwardly to scoop in the required air for the ramjet combustion system. Unfortunately, such inlets increase the complexity of the rocket and their presence could adversely affect cooling of the ramjet combustion chamber by disturbing the airflow over the outer surface of the rocket downstream of the inlets.

Furthermore, during manoeuvring of the rocket, the inlet conditions at each of the inlets varies which complicates the inlet flow to the combustor.

In another rocket type a single inlet mounted near the nose of the rocket passes air via a duct to the rear of the rocket to support the combustion of fuel at a combustor positioned ahead of the rocket fuel propellant chamber.

In both of the foregoing rocket arrangements combustion is initiated at a flameholder positioned downstream of a diffusing passage. In this diffusing passage air received from the inlet(s) undergoes firstly a supersonic diffusion process and then subsonic diffusion process. The combustion process proceeds downstream of the tlameholder within the rocket fuel propellant chamber before subsequent discharge from the rocket via a propulsion nozzle at the downstream end of this chamber.

The present invention seeks to provide a rocket which results in a more efficient use of the space available within the rocket for carrying fuel or other payloads.

According to the present invention there is provided a rocket having a propulsion unit comprising a chamber for a rocket fuel propellant and separated from a diffusing passage communicating with an air inlet at the side of the rocket by a diaphragm, the diaphragm being collapsible following use of the rocket fuel engine to allow air to flow through said diffusing passage to support combustion of fuel supplied to a flame holder and at least a part of the flame holder being normally stowed within said passage and deployable into said chamber following collapse of the diaphragm.

This arrangement saves the space normally occupied by die flame holder.

In one embodiment of the invention the flame holder comprises a pilot combustor including a fuel discharge nozzle fixedly disposed at the downstream end of the diffusing passage and a plurality of members pivotally connected around the periphery of the diffusing passage and deployable about said pivots from a stowed position in which they are radially inwardly directed to an operative position in which they extend into the said chamber, the members being disposed and shaped to support combustion of fuel and air flowing through the chamber and ignited from the said pilot combustor.

Embodiments of the invention will now be described by way of example only and with reference to the accompanying drawings in which: Fig. I is a plan view of a missile, Fig. 2 is a vertical section through the missile of Fig. I, Fig. 3 is an enlarged view of part of Fig. 2, Fig. 4 is an enlarged view of a section on the line IV-IV of Fig. 3, showing details of a collapsible diaphragm, Fig. 5 is a view in the direction V-V of Fig. 4, Fig. 6 shows details of a flame holder construction for the missile of Figs. I to 5, Fig. 7 shows an isometric view of part of the flame holder of Fig. 6.

Fig. 8 is a view of the flame holder of Fig. 6 on the line VIII-VIII.

In Figs. I and 2 there is shown a rocket missile 10 which normally flies in the attitude illustrated in Fig. 2 and which is provided with three wings 11. 12. 13 at its middle for providing lift and with further canard wings 14, 15. 16 at its forward end for controlling the attitude of the missile. Both the canard wings and the lifting wings are of curved profile and pivoted about axes 17, 18 which are fore and aft of the missile.

This allows the wings to be folded flush with the missile's cylindrical surface during launching, such as from a submarine torpedo tube. and to be deployed during tlight of the missile. At the forward portion 19 of the missile are located a guidance and control module 21, a warhead 22 and a liquid fuel tank 23. The missile is provided at its rear end with a propulsion assembly 24 which can be seen in more detail by referring also to Fig. 3.

The propulsion assembly 24 comprises a solid propellant rocket boost unit 25, which is used to provide the initial acceleration of the missile and which is subsequently detached from the missile nozzle 26 by detonating a circumferential array of explosive bolts 27. Upstream of the nozzle 26 is a chamber 28 whose outer wall 29 is a cylinder made from helically wound and welded maraging steel strip and which forms part of the missile casting. The chamber 28 is filled with a solid fuel rocket propellant 30 and lined with an ablative lining 31. The ablative lining is of a type well known per se and could for example, comprise either a silicone elastomer such as Dow Coming 93 - 104 or a matrix of silica fibres impregnated with phenolic resin such as Phenolic Refrasil solid by The chemical Insulating Co. Ltd.

After detachment of the boost rocket the missile control system initiates the ignition of the solid fuel rocket propellant which progressively bums and develops, via the discharge nozzle 26, a propulsive thrust for accelerating the missile to speeds in excess of Mach 2.0. The ablative lining 31 gradually decomposes under the effects of the burning propellant but lasts sufficiently long to prevent the buming propellant from deieteriously affecting the chamber wall. The nozzle throat 32 is fitted with an ablative liner upstream of the nozzle throat and the throat itself is fitted with a molybdenum liner 33 to withstand erosion by the exhaust gas.

The pressure generated by the buming propellant is reacted by a generally convex pressure diaphragm 34 whose detail construction can be more readily seen from Figs. 4 and 5.

The diaphragm comprises basically a circumferential array of titanium sectors 35 which abut together to form a convex cone of large cone angle and which is closed by a central titanium plug 36. The cone and plug are overlaid on the pressure surface with a single layer of glass fibre reinforced plastic 37 to ensure the sectors remain correctly aligned and abutted during assembly. The fibre reinforced plastic 37 is protected by a further layer of ablative lining 38. Whilst the fibre reinforced plastic layer 37 is shown as being attached to the convex surface of the diaphragm 34 adjacent the solid fuel rocket propellant, it could alternatively be attached to the concave surface of the diaphragm 34. In this case the ablative lining is placed directly against the convex surface of the diaphragm 34.

Upstream of the pressure diaphragm 34 is a passage 39 communicating with a chin type inlet 40 adjacent a waisted portion 41 of the missile casing. The inlet 40 is normally closed by a shield 42 which smoothly fairs the inlet into the cylindrical contour of the missile. The shield 42 is blown away by explosive retaining bolts not shown but well-known per se as soon as the rocket fuel propellant has completed its buming.

Ram air then pressurises the passage 39 and the pressure fractures the diaphragm into disposable pieces which are ejected through the discharge nozzle 26. The subsequent flow of air through the passage 39, which is shaped as a diffuser, supports the combustion of further fuel supplied to a flame holder illustrated generally at 60 and shown disposed in a stowed inoperative position within the passage 39. As will be later explained with reference to Figs. 6. 7. 8 the flame holder is deployed into the rocket fuel propellant chamber following collapse of the diaphragm 34. This further fuel is stored in the tank 23 adjacent the waisted portion of the rocket. The combustion process continues in the chamber 28 which is vacant after the buming of the solid fuel rocket propellant.The discharge nozzle 45 for the products of the ramjet combustion process is made operative after burning of the solid rocket fuel propellant by detonating a further ring of explosive bolts 46 which releases the rocket discharge nozzle 32. The ramjet discharge nozzle 45 is cooled by virtue of free stream air flowing through apertures 47 in the missile casing, impringing on film cooling the outer surface of the nozzle 45. Flow through the discharge nozzle 45 entrains the film cooling flow via orifices 48.

The steel outer wall of the combustion chamber is cooled during this ramjet combustion process, which can be of considerable duration, by the free stream of air flowing over the surface of the missile. For this cooling to be effective it is important that there are no structures upstream of the chamber which adversely disturb the airflow over the outer surface of the chamber wall. It is also important that any cable runs that have been utilised in connection with firing of the boost rocket, vectoring of the thrust nozzle of the boost rocket, separation of the boost rocket from the missile or ignition of the solid fuel rocket propellant are housed so that they do not interfere with the cooling of or heat transfer across the chamber wall. This is conveniently achieved by utilising thin cables secured with adhesive to the outer casing of the rocket or alternatively running in longitudinal grooves in the missile casing if it is to be launched from a closely fitting launching tube.

The cables are tom away at the time of ignition of the ramjet combustion system on release of the discharge nozzle 32. This is not a problem because after this time the cables are no longer required.

Referring now also to Figs. 6. 7. 8 the flame holder 60 can be seen in more detail in both its stowed and its operative position. The flame holder 60 has a small fixed apertured conical portion 61 which forms a pilot combustor and includes a fuel discharge nozzle centrally located in the passage 39 and a radial array of members 62. The members are provided with pivots at their radially outer ends and after buming of the solid fuel rocket propellant and following disposal of the diaphragm the fingers swing outwardly into the position 63. In this position small projections 64 on the members abut on corresponding rim portions 65 of the chamber 28 so that the members lie directed in a downstream direction forming longerons for the flame holder.

The members, as can be seen from the isometric view of Fig. 7 are of T section and co-operate with the rim ponions 65 to establish vortices for supporting the combustion of fuel and air flowing through the chamber and ignited from pilot combustion within the apertured cone 61.

It will be appreciated that many other forms of flame holders can be conceived which are deployable from a stowed inoperative position within the diffusion passage to an operative position in which they extend panially or wholly into the rocket fuel propellant chamber.

It will be appreciated that several modifications could be incorporated into the missile described. In particular the boost rocket can be regarded as an auxiliary feature not necessary for a missile of shorter range and the solid fuel rocket propellant cold be replaced by a liquid fuel rocket propellant. The missile illustrated has the intake adjacent a central waisted portion so as to facilitate launch from a launching tube but there is no reason why, for other means of launching. the inlet should not be placed alongside a generally cylindrical missile casing.

It will be particularly understood that whilst the above arrangement relates to a missile there is no reason why the warhead should not be replaced with reconnaissance equipment or the rocket used as a moving target.

WHAT WE CLAIM IS: 1. A rocket having a propulsion unit comprising a chamber for a rocket fuel propellant and separated from a diffusing passage communicating with an air inlet at the side of the rocket by a diaphragm. the diaphragm being collapsible following use of the rocket fuel engine to allow air to flow through said diffusing passage to support combustion of fuel supplied tc a flame holder and at least a part of the flame holder being normally stowed within said passage and deployable into said chamber following collapse of the diaphragm.

2. A rocket according to claim I characterised in that the flame holder comprises a pilot combustor including a fuel discharge nozzle fixedly disposed at the downstream end of the diffusing passage and a plurality of members pivotally connected around the periphery of the diffusing passage and deployable about said pivots from a stowed position in which they are readily inwardly directed to an operative position in which they extend into the said chamber. the members being disposed and shaped to support combustion of fuel and air flowing through the chamber and ignited from said pilot combustor.

3. A rocket according to claim 7 and characterised in that the members are provided with means for abuttingly engaging the wall of the diffusing passage to control, in their operative position. their inclination with respect to the said chamber.

4. A rocket according to claim 2 and characterised in that the members are of tee section and are disposed in their operative position with the vertical leg of the tee facing radically inwardly into the said chamber.

5. A rocket according to claim 1 and characterised in that the collapsible diaphragm comprises an array of generally triangular segments closely fitted together to define a shallow angle cone with a central aperture at the apex of the cone and having a plug to close the aperture. the cone being overlaid on its outer surface with a layer of fibre reinforced resin and a further layer of an ablative coating.

6. A rocket substantially as herein described and illustrated with reference to the accompanying drawings.

**WARNING** end of DESC field may overlap start of CLMS **.

Claims (6)

**WARNING** start of CLMS field may overlap end of DESC **. is to be launched from a closely fitting launching tube. The cables are tom away at the time of ignition of the ramjet combustion system on release of the discharge nozzle 32. This is not a problem because after this time the cables are no longer required. Referring now also to Figs. 6. 7. 8 the flame holder 60 can be seen in more detail in both its stowed and its operative position. The flame holder 60 has a small fixed apertured conical portion 61 which forms a pilot combustor and includes a fuel discharge nozzle centrally located in the passage 39 and a radial array of members 62. The members are provided with pivots at their radially outer ends and after buming of the solid fuel rocket propellant and following disposal of the diaphragm the fingers swing outwardly into the position 63. In this position small projections 64 on the members abut on corresponding rim portions 65 of the chamber 28 so that the members lie directed in a downstream direction forming longerons for the flame holder. The members, as can be seen from the isometric view of Fig. 7 are of T section and co-operate with the rim ponions 65 to establish vortices for supporting the combustion of fuel and air flowing through the chamber and ignited from pilot combustion within the apertured cone 61. It will be appreciated that many other forms of flame holders can be conceived which are deployable from a stowed inoperative position within the diffusion passage to an operative position in which they extend panially or wholly into the rocket fuel propellant chamber. It will be appreciated that several modifications could be incorporated into the missile described. In particular the boost rocket can be regarded as an auxiliary feature not necessary for a missile of shorter range and the solid fuel rocket propellant cold be replaced by a liquid fuel rocket propellant. The missile illustrated has the intake adjacent a central waisted portion so as to facilitate launch from a launching tube but there is no reason why, for other means of launching. the inlet should not be placed alongside a generally cylindrical missile casing. It will be particularly understood that whilst the above arrangement relates to a missile there is no reason why the warhead should not be replaced with reconnaissance equipment or the rocket used as a moving target. WHAT WE CLAIM IS:

1. A rocket having a propulsion unit comprising a chamber for a rocket fuel propellant and separated from a diffusing passage communicating with an air inlet at the side of the rocket by a diaphragm. the diaphragm being collapsible following use of the rocket fuel engine to allow air to flow through said diffusing passage to support combustion of fuel supplied tc a flame holder and at least a part of the flame holder being normally stowed within said passage and deployable into said chamber following collapse of the diaphragm.

2. A rocket according to claim I characterised in that the flame holder comprises a pilot combustor including a fuel discharge nozzle fixedly disposed at the downstream end of the diffusing passage and a plurality of members pivotally connected around the periphery of the diffusing passage and deployable about said pivots from a stowed position in which they are readily inwardly directed to an operative position in which they extend into the said chamber. the members being disposed and shaped to support combustion of fuel and air flowing through the chamber and ignited from said pilot combustor.

3. A rocket according to claim 7 and characterised in that the members are provided with means for abuttingly engaging the wall of the diffusing passage to control, in their operative position. their inclination with respect to the said chamber.

4. A rocket according to claim 2 and characterised in that the members are of tee section and are disposed in their operative position with the vertical leg of the tee facing radically inwardly into the said chamber.

5. A rocket according to claim 1 and characterised in that the collapsible diaphragm comprises an array of generally triangular segments closely fitted together to define a shallow angle cone with a central aperture at the apex of the cone and having a plug to close the aperture. the cone being overlaid on its outer surface with a layer of fibre reinforced resin and a further layer of an ablative coating.

6. A rocket substantially as herein described and illustrated with reference to the accompanying drawings.