JPH0245644A - Membrance seal assembly for solid propellant rocket engine - Google Patents

Abstract

PURPOSE: To enable a combustion chamber to be separate by arranging a thin tray-shaped diaphragm made of ductile material of high strength on the front end side of a first case in a hole closing relationship. CONSTITUTION: A thin tray-shaped diaphragm 100 made of ductile material of high strength is arranged on the front end side of a first case 86 in a relationship of closing a hole 98 of a cylindrical structure support member 90. An insulation material layer 102 is bonded to the side of the diaphragm 100, facing a combustion chamber 104. Thus, it is possible to form a high pressure on the diaphragm 100 side of the structure support member 90, which is higher than that on the other side of the structure support member 90, and it is possible to separate the combustion chamber.

Description

DETAILED DESCRIPTION OF THE INVENTION [Industrial Application Field] The present invention relates to an improved solid propellant rocket engine or gas generator, and more particularly, to an improved solid propellant rocket engine or gas generator having a plurality of solid propellant units and having a plurality of solid propellant units and which can be controlled by commands. The present invention relates to a solid propellant rocket engine equipped with a membrane seal structure that allows ignition to be performed independently of each other to obtain differentiated propulsive forces.

[Prior art and problems with the invention]

The entire propulsion capacity of a solid propellant rocket engine is normally consumed during the combustion process of one solid propellant weight. This is because once the solid propellant is ignited, it is very difficult to stop the combustion process until the entire weight of the solid propellant has been consumed.

In the prior art, there have been proposals to provide - solid propellant rocket engines that can be ignited in a controlled manner, i.e. start-stop-restart rocket engines, for example two or more concentric solid propellants; This has been accomplished by providing zones of multiple layers of solid propellant within the propellant unit, ie, the combustion chamber, and separating the layers of solid propellant by combustion inhibiting barriers. This barrier confined combustion to a single layer of solid propellant and was then made of a material that was susceptible to rupture for ignition of an adjacent layer of solid propellant.

One such prior art for a rocket engine that can be ignited more than once is disclosed in U.S. Pat.
No. 3.855, wherein a pyrotechnic material and an electrically ignitable film are provided between each of the layers for commanded ignition, and then between each of the adjacent layers. .

Other such prior art patents include U.S. Pat. Patent No. 3.56
No. 8.448, one of two concentric layers of solid propellant separated by a combustion inhibit barrier is adapted to be ignited by an ignition device extending from the nozzle of the rocket engine into the combustion chamber. The other layer is ignited by a gas generator connected to the head end of the combustion chamber by a tubular extension. A frangible membrane seal and perforated support member assembly is provided to isolate the gas generator from the combustion chamber of the engine during combustion of the first rocket propellant layer.

U.S. Pat.
U.S. Pat. No. 3,354,647, issued Jan. 28 to Double Sea Icock, is of a similar construction, but it uses a combustible barrier for destruction of the combustible barrier and ignition of the adjacent propellant. Discloses introducing liquid fuel into the chamber.

All such prior art patents include multiple concentric layers or zones of solid propellant in a single combustion chamber, which then initiate combustion and ignition of adjacent layers of propellant. involving the destruction of a burn inhibit barrier disposed between the layers in order to ignite the propellant; It is. The Breedman Jr. et al. and Webb Jr. patents further encompass the use of breakable membranes and perforated support members, which tend to introduce debris into the combustion chamber when the membranes are broken and disassembled, respectively. be. The Mangan and Aycock patent further includes the introduction of liquid fuel into the combustion chamber.

It is an object of the present invention to be able to be ignited more than once, i.e. in a pulsed manner, and without involving the destruction of burn-inhibiting burrs'1, and without using destructible membranes or support members that can be destroyed or disassembled. , and to provide a solid propellant rocket engine and gas generator that allows the use of conventional ignition devices to ignite different solid propellant units.

A more particular object of the present invention is to provide an improvement in such solid propellant rocket engines and gas generators that comprises physically separating a plurality of combustion chambers to facilitate the production of segmented propulsion segments on command. The purpose is to provide

A further particular object of the invention is to provide such an improved solid propellant rocket engine and gas engine which can be manufactured with a number of concentric or tandemly arranged combustion chamber volumes as desired. is to provide a generator.

[Means to solve the problem]

To achieve the above object, according to the invention, thin membranes made of a high strength and ductile material, for example 0.010 inch thick nickel, are formed in a combustion chamber in a concentric or tandem arrangement. solid material applied to a perforated structural support member or bulkhead separating the combustion chambers, thereby applying a substantially higher pressure on the membrane side of the perforated structural support member than on the other side to separate the combustion chambers. A propellant rocket engine and a gas generator are provided. Applying pressure in the opposite direction to the perforated structural support member causes the membrane to collapse allowing communication between the combustion chambers.

The solid propellant rocket engine and gas generator according to the invention can be arranged in multiple concentric or tandem configurations as desired.
That is, it can be integrated into combustion chamber volumes arranged in series. Each combustion chamber volume can include one or more solid propellant units or layers.

〔Example〕

Embodiments of the present invention will be described in detail below with reference to the drawings.

Referring to FIG. 1, reference numeral 10 indicates a solid propellant rocket engine as a whole. The solid propellant rocket engine 10 includes a rocket engine case 12 made of aluminum, and a converging-dipurging type nozzle 14 is attached to the rear end of the case 12. The nozzle 14 can be formed as a phenolic molded article, and a generally cylindrical member 16 is formed integrally with the nozzle 14. This cylindrical member 16 has a reduced diameter annular structural support 18 located at the front. The cylindrical member 16 is held at the rear end of the case 12 between an inner peripheral rib 20 provided within the case and an aluminum retainer ring 22 provided at the end of the case.

A bead 24 on the circumference of the retainer ring 22 engages a circumferential groove 26 on the inner surface of the case 12 to hold the nozzle 14 and cylindrical member 16 in place, and an O-ring seal 28 and the inner surface of the case 12.

As shown in FIG. 1, a space within the case 12 defined between the forward end of the case 12 and the annular structural support 18 is defined by a tubular member 34 into first and second concentric volumes, i.e., combustion chambers. Divided into chambers 30.32. Tubular member 34 may be made of a glass phenolic material and its forward end is supported within an internal cylindrical recessed portion 36 of case 12 and held in place therein. The rearward end of the tubular member 34 is fit-supported within the annular support 18 against a shoulder 38 on the annular support 18 .

Communication between the first and second concentric combustion chambers 30.32 is provided by holes 40 provided in the annular structural support 18. The holes 40 may typically be 4 to 12 in number, with 4 holes extending from the vertical toward the rear of the rocket engine IO.
It is sloped at an angle of 5 degrees.

Hole 40 is permanently covered by a thin non-porous annular metal membrane or cover 42 made of high strength, ductile nickel, preferably o, oio inches (0,0254 cm) thick. Membrane 42 is held in place by an integrally formed flange 44 that fits between the end of tubular member 34 and shoulder 38. The first and second concentric combustion chambers 30.32 are physically separated when the outer surface of the membrane 42 is in place closing the bore 40 at all times. That is, the membrane 42 separates the first combustion chamber 30 from the second combustion chamber 32 by permitting high pressure to be applied to the membrane side of the annular structure support 18 . When pressure is applied from the second combustion chamber 32 through the hole 40 in the annular structure support 18, the membrane 42 breaks allowing communication between the first and second combustion chambers 30, 32.

A first generally cylindrical unit or layer 46 of solid propellant branes is cast or bonded to the inner wall of the case 12 between the rib 20 in the second combustion chamber 32 and the forward end of the case 12 . As shown in the drawings, the shape of the front portion of the first solid propellant unit 46 matches the curved surface of the front interior of the case 12.

The first solid propellant unit 46 is connected to the recessed portion 36 of the case.
terminating slightly short of the to provide an annular space 48, the purpose of which will be explained later. A second generally cylindrical unit or layer 50 of solid propellant branes is disposed concentrically with the first solid propellant unit 46 within the second combustion chamber 32;
This is glued to the outer surface of tubular member 34.

The forward end of tubular member 34 is closed by a disc-shaped member or plate 52 made of a suitable metal as shown. Rearly extending tubular projections 54,56 are integrally formed with the plate 52 and are bored in the plate 52 in alignment with each tubular projection 54,56 for communication.

As shown in the drawings, each tubular projection 54, 56 has a stick-type solid propellant unit 58, 6, respectively.
0 is properly attached. These solid propellant units 58 , 60 are hollow and extend rearwardly along the entire length of the tubular member 34 .
No support members are provided to support the aft ends of these solid propellant units 58,60, other than that they support each other and against the outer wall of the tubular member 34. Although only two such solid propellant units 58,60 are shown in FIG. 1, additional such solid propellant units and associated holes in plate 52 may be provided as desired so that additional It should be understood by those skilled in the art that the use of stick type solid propellant units.

In the opening formed by the recessed portion 36 between the plate 52 and the forward-most end of the case 12, first and second ignition devices 62,64 are arranged.

The first igniter 62 includes an electric squib 66 and a pyrotechnic material 68.
are placed in a cylindrical ceramic plug 70 provided with several openings. The plug 70 is provided to close the opening of the recessed portion 36 at the front end of the case 12, and is sealed by the seal ring and the thread of the plug 70 that engages with the internal thread of the opening of the recessed portion 36. Ru. Pyrotechnic material 68 is located between the inner end of plug 70 and perforated plate 52 and is slightly spaced apart by suitable spacer rings 72. A short passage 74 is provided between the electric squib 66 and the carpentry material 68. Communication between the pyrotechnic material 68 and the first combustion chamber 30 is provided by the holes in the plate 52 and tubular projections g54, 56 previously described.

The second igniter 64 includes an electric squib 76 and a pyrotechnic material 78.
, which are also arranged within the plug 70 such that the end of the electric squib 76 contacts the pyrotechnic material 78 and the metal plate 8G separates and isolates the pyrotechnic material 78 from the pyrotechnic material 68. ing. A passage 82 in the recessed portion 35, preferably a plurality of such passages, connect the pyrotechnic material 78 and the second combustion chamber 32.
provide communication between

Pyrotechnic materials 68.78 Any of a number of pyrotechnic materials can be used, typical materials being 25
It is a granular mixture of 75 weight percent boron and 75 weight percent potassium nitrate.

In operation of the rocket engine 10, the °C1 electric squib 66 is first ignited. A suitable interlock circuit is provided in the drive circuit for the electric squib 66.76 to inhibit initial firing of the electric squib 76. When the electric squib 66 is ignited, a combustion flame traverses the passageway 74 and ignites the pyrotechnic material 68. The ignition of the pyrotechnic material 68 flows through the annular space 48, the holes in the plate 52, the tubular protrusions 54, 56 to the stick-type solid propellant unit 58, 60, thereby igniting the solid propellant unit 58, 60. A high pressure is immediately created in the first combustion chamber 30 and a flow of gases propelling the rocket flows out of the nozzle 14 . The high pressure applied to the membrane side of the annular structural support 18 causes the membrane 42 to press firmly against the opening of the hole 40, thus sealing the hole 40 and separating the first combustion chamber 30 from the second combustion chamber 32. do. Combustion termination of stick-type solid propellant units 58, 60, and other similar units if provided, typically occurs very quickly. At the end of such combustion, the first
The pressure in the combustion chamber 30 of will drop very quickly, and the flow of gas out of the nozzle 14 will also drop very quickly. However, activation of the first and second solid propellant units 46,50 of the second combustion chamber 32 does not occur until the electric squib 76 is ignited.

When the electric squib 76 is ignited to ignite the pyrotechnic material 78, the combustion gases traverse the passageway 82 and ignite the first and second solid propellant units 46,50. High pressure rapidly spreads to the second combustion chamber 32.

When this high pressure is applied through the hole 40 in the annular structure support 18, the membrane 42 breaks and transfers the second combustion chamber 32 into the first combustion chamber 3.
0, thereby effecting the exit of a second gas through the nozzle 14 which propels the rocket engine.

In this way, segmented propulsion segments are obtained on command from the rocket engine 10. It will be apparent to those skilled in the art that more than one propulsion segment can be obtained by adding similar concentric solid propellant units.

2 to 4, another embodiment of the invention is shown in which the solid propellants are arranged in tandem, again allowing the solid propellant rocket engine to obtain segmented propulsion segments on command. It looks like this.

The solid propellant rocket engine 84 shown in FIG. 2 includes a first rocket engine case 86 having a plast tube 88 attached to its aft end. The illustrated plast tube 88 combines the functions of a converging-damping type nozzle, a plast tube, and a suitable flange for attachment to the case 86 in a single unit. A perforated, generally cylindrical structural support member 90 is attached to the forward end of the first case 86 to close the forward end of the first case 86, and the cylindrical structural support member 90 in turn connects to the second rocket engine case. 92, and the ends of the first and second cases 86.92 are placed in overlapping relationship. The front and rear ends of the cylindrical structural support member 90 are dish-shaped. A hole 98 extending parallel to the axis of rocket engine 84 is provided in cylindrical structural support member 90 .

The first and second cases 86.92 and the plast tube 88 are made of any suitable material, and a suitable phenolic molded insulation layer is provided on the first and second cases 86.92 and the plast tube 88. . A solid propellant unit or layer 94 is provided within the first case 86 and a similar solid propellant unit or layer 96 is provided within the second case 92.

A thin dish-shaped membrane or cover 100 of high strength, ductile material, such as nickel, is disposed on the front end side of the first case 86 in closing relation to the hole 98 in the cylindrical structural support member 90. corresponds to the shape of the rear end side of the cylindrical structural support member 90. Soft insulating material layer 10 such as polyisoprene
2 is formed within the first case 86 of the membrane 100.
It is glued to the side facing the combustion chamber 104 and is turned back there. A second combustion chamber 106 is formed within second case 92 .

FIG. 3 shows an enlarged view of the cylindrical structural support member 90, the membrane 100, and the polyisoprene insulation layer 1020.

In order to ignite the first solid propellant unit 94, the first
An igniter 108 is attached to the nozzle and plast tube 88. A second igniter 110 is suitably attached to the front end of the second case 92 to ignite the second solid propellant unit 9G.

As shown in FIG. 4, the membrane 100 is preferably 0,0
10 to 0.020 inches (0.0254 to 0.05
Made as triangular or pie-shaped stacked pieces of 08 cm) nickel Zino stock. FIG. 4 shows a plurality of stacked pie-shaped strips in effective relation to close the hole 98 on the side of the cylindrical structural support member 90 facing the combustion chamber 104.

In operation of a rocket engine arranged in tandem, the ignition of the first igniter 108 is caused by the ignition of the first solid propellant unit 94.
ignition, and high pressure is rapidly created in the combustion chamber 104. The high pressure on the membrane 100 side of the cylindrical structural support member 90 forces the polyisoprene insulation layer 102 and membrane 100 firmly against the hole 98 in the cylindrical structural support member 90, thus sealing the hole 98 and opening the first combustion chamber 104 from a second combustion chamber 106. As a result, the flow of gas that propels the rocket flows through the nozzle and plast tube 88.
flows out from

When the first solid propellant unit 94 finishes burning, the first solid propellant unit 94
The pressure in the combustion chamber 104 decreases, and the flow of gas exiting the nozzle and plast tube 88 decreases as well.

Ignition of the second solid propellant unit 96 of the second combustion chamber 106 does not occur until the second igniter 110 is ignited.

When the second ignition device 110 is ignited, the interlock that prohibited the second ignition device 110 from being ignited until the first ignition device 108 is ignited is removed, and the second combustion chamber 106 High pressure builds up. This high pressure ruptures the membrane 100 and the center of the pie-shaped piece of membrane 100 is forced outward as shown in FIG. 5, rupturing the polyisoprene insulation layer 102 to which it was adhered. This allows the holes 98 to pass through the first and second combustion chambers 104,1.
06 is in communication and a second flow of gas that propels the rocket exits the nozzle and plast tube 88.

The segmented propulsion segments are achieved by commands from the tandemly arranged rocket engines 84. As will be understood by those skilled in the art, additional tandem or series volumes may be provided as desired to provide additional segmented propulsion segments.

Accordingly, in accordance with the present invention, high strength, ductile thin nickel films ranging in thickness from 0.010 to 0.020 inches (0.0254 to 0.0508 cm) are fabricated in a concentric or tandem arrangement. It is applied to the perforated structural support means that separates the combustion chamber of the rocket engine, so that a higher pressure can be created on the ventral side of the structural support means than on the other side of the structural support means, which separates the combustion chamber. A membrane seal assembly for use in a pulsed rocket engine is provided. If pressure is applied in the opposite direction to the perforated structural support means, the membrane will rupture and the two combustion chambers will be in communication.

[Brief explanation of the drawing]

FIG. 1 is a partial cross-sectional side view showing that a membrane seal structure according to the present invention physically separates the concentric combustion chambers of a rocket engine, and FIG. 2 is a membrane seal of another embodiment according to the present invention. a partial cross-sectional side view showing the structure physically separating the tandem combustion chambers of a rocket engine;
Figure 3 is a partially enlarged view of the structural support member and membrane in Figure 2, Figure 4 is a front view of the membrane in Figure 2, and Figure 5 shows the membranes in Figures 2 to 4 breaking. FIG. 12... Case, 14... Nozzle, 18
... Structural support part, 30.32 ... Combustion chamber, 40
...hole, 42...membrane, 46, 50.5
8.60... Solid propellant, 86.92... Case, 88... Nozzle and plast tube, 90... Structural support member, 94.96... Solid propellant, 98...
・Hole, 100...Membrane, 104.106
...combustion chamber. Procedural Amendment April 2, 1999 Commissioner of the Japan Patent Office Yoshi 1) Mr. Moon Takeshi! , Indication of the case Patent Application No. 194694 of 1988 2, Title of the invention 5 for solid propellant rocket engine, Column 6 of "Claims" of the specification subject to amendment, Contents of the amendment Scope of claims Correct as shown in the attached sheet. 7. List of attached documents Claims 1 copy 4. Agent address: Shizuko Toranomon Building, 8-1o Toranomon-chome, Minato-ku, Tokyo 105 Phone: 504-0721 Name: Patent attorney (65
79) Akira Aoki. (4 others) "' 2. Claim 1: A membrane seal assembly for a solid propellant rocket engine having at least two solid propellant units, the membrane seal assembly comprising: separating said solid propellant units; a rocket engine comprising structural support means defining a plurality of combustion chambers each containing the solid propellant unit, the structural support means having holes for providing communication between the plurality of combustion chambers; a membrane seal made of a strong, ductile material and disposed over the hole in one of the combustion chambers, the membrane seal providing a substantially higher pressure in the one combustion chamber than in the other combustion chamber; prohibiting communication between the combustion chambers when the other combustion chamber has a substantially higher pressure than the one combustion chamber; A membrane seal assembly in which the seal is dish-shaped and made of pre-shaped shim stock pieces arranged one above the other. 2. A solid propellant rocket engine having at least two solid propellant units. a membrane seal assembly for: a membrane seal assembly for separating said solid propellant units to form a plurality of combustion chambers in said rocket engine, each containing said solid propellant units; having a sump hole providing communication between the plurality of combustion chambers, and having a diameter of 0.010 to 0.020 inches
．． a membrane seal made of a high-strength, ductile material having a thickness in the range of 0.0508 cm) and disposed over the hole in one of the combustion chambers; prohibiting communication between said combustion chambers when there is a substantially higher pressure than said other combustion chamber; and said when said other combustion chamber has a substantially higher pressure than said one combustion chamber. A membrane seal assembly allowing communication between the combustion chambers, the membrane seal being dished and made of pre-shaped pieces of nickel shim stock and placed one above the other. 3. A membrane seal assembly for a solid propellant rocket engine having at least two solid propellant units, the solid propellant units being separated and a plurality of solid propellant units each housed in the rocket engine. a structural support means forming a combustion chamber of the plurality of combustion chambers, the structural support means having holes for providing communication between the plurality of combustion chambers, the structural support means being made of a high strength, ductile material and one of the combustion chambers; a membrane seal disposed over the hole, the membrane seal prohibiting communication between the combustion chambers when there is a substantially higher pressure in the one combustion chamber than the other combustion chamber; and allowing communication between the combustion chambers when the other combustion chamber has a substantially higher pressure than the one combustion chamber, the membrane seal is nonporous, and the structural support means is annular. forming and concentrically arranged separate combustion chambers, said one combustion chamber being an inner combustion chamber, and said membrane seal being annularly formed so that its outer surface is always in contact with said structural support means. a membrane seal assembly disposed to close said hole; 4. A membrane seal assembly for a solid propellant rocket engine having at least two solid propellant units, wherein the solid propellant units are separated and a plurality of solid propellant units are each housed in the rocket engine. a structural support means defining a plurality of combustion chambers, the structural support means having holes for providing communication between the plurality of combustion chambers;
．． a membrane seal made of a high-strength, ductile material having a thickness in the range of 0.0508 cm) and disposed over the hole in one of the combustion chambers; prohibiting communication between said combustion chambers when there is a substantially higher pressure in said other combustion chamber than said one combustion chamber; allowing communication between the combustion chambers, the membrane seal being made of nickel and being non-porous, the structural support means being annularly shaped and forming separate combustion chambers arranged concentrically;
The one combustion chamber is an inner combustion chamber, the hole is inclined at a predetermined angle from vertical toward the rear of the rocket engine, and the membrane seal is formed in an annular shape such that its outer surface is always connected to the structural support means. a membrane seal assembly positioned to close said hole of the membrane.

Claims (1)

Claims: 1. A membrane seal assembly for a solid propellant rocket engine having at least two solid propellant units, the membrane seal assembly comprising: separating the solid propellant units and attaching the solid propellant units to the rocket engine; a structural support means defining a plurality of combustion chambers each containing a combustion chamber, the structural support means having holes for providing communication between the plurality of combustion chambers, the structural support means having holes between 0.010 and 0.020 inches ( 0.0254 to 0
．． a membrane seal made of a high-strength, ductile material having a thickness in the range of 0.0508 cm) and disposed over the hole in one of the combustion chambers; prohibiting communication between said combustion chambers when there is a substantially higher pressure than said other combustion chamber; and said when said other combustion chamber has a substantially higher pressure than said one combustion chamber. A membrane seal assembly allowing communication between the combustion chambers. 2. The membrane seal assembly of claim 1, wherein the membrane seal is made of nickel. 3. The membrane seal assembly of claim 2, wherein the membrane seal is non-porous. 4. The membrane seal assembly of claim 2, wherein said membrane seal is dish-shaped and made of stacked pre-shaped pieces of nickel shim stock. 5. The structural support means is annularly shaped and forms separate combustion chambers arranged concentrically, the one combustion chamber being an inner combustion chamber, and the hole facing toward the rear of the rocket engine. 4. A membrane seal assembly according to claim 3, wherein said membrane seal is annularly formed and positioned such that its outer surface always closes said hole in said structural support means. 6. A claim comprising: a first ignition means for igniting the solid propellant unit in the one combustion chamber; and a second ignition means for igniting the solid propellant unit in the other combustion chamber. The membrane seal assembly according to item 5. 7. said structural support means being cylindrical with a rearwardly facing dish-shaped surface and forming separate combustion chambers in a tandem arrangement;
the one combustion chamber is aft of the other combustion chamber, the hole extends parallel to the axis of the tube of the structural support means, and the membrane seal is dish-shaped to match the dish-shaped surface of the structural support means. 5. A membrane seal assembly as claimed in claim 4, formed and arranged to close said hole in said structural support means at all times.