US3316718A - Honeycomb structured propellant for rocket motors - Google Patents

Honeycomb structured propellant for rocket motors Download PDF

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US3316718A
US3316718A US425354A US42535465A US3316718A US 3316718 A US3316718 A US 3316718A US 425354 A US425354 A US 425354A US 42535465 A US42535465 A US 42535465A US 3316718 A US3316718 A US 3316718A
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burning
grain
propellant
honeycomb
cell
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US425354A
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Wells A Webb
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Lockheed Corp
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Lockheed Corp
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K9/00Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08Rocket- engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/10Shape or structure of solid propellant charges
    • F02K9/16Shape or structure of solid propellant charges of honeycomb structure
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00Geometry
    • F05D2250/20Three-dimensional
    • F05D2250/28Three-dimensional patterned
    • F05D2250/283Three-dimensional patterned honeycomb
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00Geometry
    • F05D2250/60Structure; Surface texture

Description

W. A. WEBB May 2, 1967 HONEYGOMB STRUCTURED PROPELLANT FOR ROCKET MOTORS 5 Sheets-Sheet 1 Filed Jan. 13, 1965 DOG EAR INVENTOR WELLS A. WEBB W. A. WEBB May 2, 1967 HONEYCOMB STRUCTURED PROPELLANT FOR ROCKET MOTORS 3 Sheets-Sheet 2 Filed Jan. 13, 1965 FIGB INVENTOR. WELLS AWEBB W. A. WEBB May 2, 1967 HONEYCOMB STRUCTURED PROPELLANT FOR ROCKET MOTORS 3 Sheets-Sheet 5 Filed Jan. 13, 1965 FIG.5

DIRECTION OF BURNING DIRECTION OF BURNING INVENTOR. WELLS A. WEBB United States Patent ()fidce 3,316,718 Patented May 2, 1967 3,316,718 HONEYCUMB STRUCTURED PROPELLANT FOR ROCKET MOTORS Wells A. Webb, Redlands, Calif., assignor to Lockheed Aircraft Corporation, Los Angeles, Calif. Filed Jan. 13, 1965, Ser. No. 425,354 11 Claims. (til. 60-255) This invention relates to propellant grains for rocket motors, and more particularly to improvements in solid propellant grains which are reinforced with a honeycomb cellular structure.

Metal honeycomb cellular structures, or cores, have been previously proposed for use in end-burning, solid propellant grains to give better mechanical properties and to increase the burning rates of the grains. However, these proposals have not been entirely successful, and it has now been determined that the ditficulties encountered were due in large measure to the cell orientation of the honeycomb core with respect to the combustion front, or burning surface. In the previous proposals, the cells were oriented so that the ribbons comprising the cell walls were perpendicular to the combustion front; i.e., the general direction of burning was in the direction of the cell axis, or T direction. Thus, if there should be a separation of the solid propellant from the cell Wall in some localized area due perhaps to shrinkage of the propellant upon cure, the flame could advance down the gap ahead of the normal combustion front and cause uneven burning or, in some cases, failure of the rocket motor. The use of viscous liquid propellants prevented this kind of failure, but introduced other problems involving the control of the liquid in the honeycomb core which still remain to be solved.

Another difiiculty of the previously proposed endburning, solid-propellant grains which burned in the T direction was a low initial chamber pressure and a relatively slow buildup of such pressure due to the fact that the ignition surface was flat and consisted primarily of solid propellant. Upon ignition, there was a time lag before the flat propellant surface in each cell could be converted into a conical, or dentated surface characteristic of the steady state burning condition.

In accordance with the present invention, the previous problems have been overcome by orienting the honeycomb core so that burning takes place in the ribbon or L direction; i.e., the cell orientation has now been turned 90 from the previous proposals. Now, the solid fuel mass within each cell is completely surrounded by the cell walls, and the advancing flame must first consume the overlying cell wall portions before the underlying solid fuel mass can be burned. Any cracks or gaps which might develop in the fuel are now confined to a single layer or row of cells, and there is no possibility that the flame can advance prematurely through the entire grain.

The present cell orientation also produces a quicker buildup of chamber pressure by providing for initial exposure of the honeycomb cell walls and dog ears. These are cleaned of propellant and then coated with a special igniter composition. Because of the higher heat conductivity of the metal surfaces initially exposed, there is very little time lag between the ignition stage and the steady state burning stage, the latter being characterized by a corrugated burning surface.

Also, as in the previous proposals, it has been found that the present mode of cell orientation gives the same advantages of reinforcement of the solid fuel and of increased burning rate. Generally, the present grains burn at about three times the rate of burning of the same propellant without the honeycomb core.

These advantages and other features of the present invention will be further described in detail in conjunction with the accompanying drawings wherein:

FIGURE 1 is a perspective, schematic view of a portion of a honeycomb core, illustrating core nomenclature and directions;

FIGURE 2 is a fragmentary, schematic, sectional view of a rocket motor containing an end burning propellant grain of the present invention;

FIGURE 3 is a fragmentary, schematic, sectional view of a rocket motor containing another end burning propellant grain of the present invention, said grain having a truncated ignition surface;

FIGURE 4 is a fragmentary, schematic, sectional view of a rocket motor containing another end burning propellant grain of the present invention;

FIGURE 5 is a sectional view of an. internal burning, segmented propellant grain of the present invention;

FIGURE 6 is a perspective view of a segment of another internal burning, segmented propellant grain of the present invention, showing the honeycomb core pieces comprising the segment in staggered relationship; and

FIGURE 7 is an end view of still another internal burning, segmented propellant grain of the present invention.

FIGURE 2 shows a rocket motor 10 with a forward or head end 12, a generally cylindrical combustion chamber, or case, 14, and an exhaust nozzle 16. Case 14 contains an end burning solid propellant grain 18 of the present invention. Grain 18 may be ignited at the aft end by any conventional means, such as an electrically-actuated squib 20, and produces combustion gases which are discharged through nozzle 16 to propel the motor forward in well known manner.

Grain 18 is prepared for loading into motor machining a honeycomb core 22, preferably aluminum, into cylindrical shape with the cell walls, or ribbons 24, running in the direction of burning and with the cells being expanded or fully expanded depending on whether propellant 26 is a soft cure material or an uncured material. Propellant 26 is cast into the cells under vacuum and then cured. A combustion inhibiting liner 28 is then applied on the cylindrical and head end surfaces and cured. Such a liner may be comprised of multiple coats of any suitable combustion inhibitor such as one containing 30% by weight epoxy adhesive, 68% polyamid, and 2% carbon black. The aft, or ignition, surface is cleaned of all propellant 26, and the exposed cell walls and dog ears 29 are coated with a suitable igniter composition 30, such as one containing parts by weight magnesium powder, 25 parts p0- tassium perchlorate, 15 parts barium nitrate, and 5 parts polyisobutylene binder in a volatile vehicle such as petroleum ether. The grain thus prepared is then loaded into the motor as a free-standing structure, with the cell axis perpendicular to the thrust axis.

The free-standing grain may be fastened to the motor by any suitable means, such as that shown in FIGURE 2 Flange 25 is adapted to be held in clamped relationship 14 and the head end 12, the latter being screwed onto the case. As is known, the free standing structure allows for a gap 27 between the case and the grain. The gap permits small movements of the grain due to thermal expansion and contraction, thereby relieving any stresses which may be set up in the grain. The gap is pressurized with static gases when the motor is fired.

Aluminum alloy 5052 or some other alloy containing from about /2% to 5% magnesium is preferred for the honeycomb core because of the higher burning efliciencies obtained. Aluminum alloys without magnesium may also be used, but the burning efficiencies are lower. The burning efficiency is also dependent upon the thickness or gauge of the honeycomb foil material. Generally, as thin a gauge as possible, consistent with strength requirements, is selected, with gauges in the range of 0.9 to 2 mils being preferred.

Propellant 26 may be any propellant having a viscosity which is now enough to permit flow into the honeycomb cells. Generally, the viscosity of such a propellant will be below 200,000 centipoises, with viscosities in the range of 40,000 to 140,000 centipoises being preferred. Typical examples of such propellants contain a solid oxidizing salt, such as ammonium perchlorate crystals, as a major component and a polymerizable organic compound, such as polysulfide, polyurethane, and polybutadiene, as a minor component.

Examples of specific propellant formulations suitable for use in the present invention are shown below:

EXAMPLE I Ingredient: Percent by weight Polybutadiene 7 Aluminum powder 13 Aluminum perchlorate 69 Polyester plasticizer 8 Epoxy resin 1 Burning rate catalyst 2 EXAMPLE II Ingredient: Percent by weight Ball powder (nitrocellulose-nitroglycerine) 15 Nitrate esters 32 Aluminum powder 17 Ammonium perchlorate 26 Ethyl centralite 0.5 Resorcinol 0.5 Cyclotetramethylene tetranitramine 9 Test strands made of the propellant shown in Example I above burned at the rate of 1 inch per second at a pressure of 1,000 pounds per square inch, producing a pressure exponent, n, equal to 0.29 in the burning rate equation:

where r is the burning rate, C is a function of the grain temperature, and p is chamber pressure. When the same propellant was cast into a honeycomb core, the burning rate increased to 3 inches per second at 1,000 pounds per square inch and the pressure exponent dropped to 0.20, showing that the honeycomb core produces an accelerated burning rate and a greater safety and stability of chamber pressure. The threefold increase in apparent burning rate is accounted for by the fact that the steady state burning surface is corrugated and has approximately three times the area of surface found in the conventional grains without honeycomb.

It has been found that the combustion time versus chamber pressure curve for grain 18 of FIGURE 2 shows regularly occurring undulations in pressure which apparently result from the burning through of overlying cell walls in successive cell layers. In applications requiring a smoother pressure curve, grain 18 may be modified as shown in FIGURE 3 by providing for an inclined, planar ignition surface 32. Here, the fact that the burning front advances at an angle to the longitudinal axis of the rocket motor assures that the overlying cell walls in the successive cell layers are burned through at different times, thus smoothing out the chamber pressure.

The same effect of smoothing out the pressure curve may be obtained by modifying the grain as shown in FIGURE 4. Here, the cylindrical core of FIGURE 2 has been divided longitudinally and the respective halves 4 34, 36 have been bonded together with the cells staggered or offset approximately one-half of the cell length. With this core structure, the burning through of successive cell layers in the two halves is out of phase with each other and a smoother pressure curve is obtained.

FIGURE 5 shows a typical end section of an internal, or central, burning propellant grain comprised of three longitudinal segments 38, each having opposing convex surfaces 42, 44 and the general configuration shown in FIGURE 6. These segments are made by fan-wise expansion of the honeycomb core 46 with the cell ribbons running generally in the direction of burning, as shown. After shaping to form the convex surface 44, segments 38 are filled with a suitable solid propellant 48 in the same manner as described for grain 18. Multiple coats of a suitable burning inhibitor 50 are then applied to surface 44 and to the end surfaces (not shown) of the grain. Surface 42 including dog ears 43 is cleaned of propellant and a suitable igniter composition 52 is applied thereto and cured. Segments 38 are then bonded longitudinally along a central portion 54 to a sleeve 56, as shown in FIGURE 5. Sleeve 56 thus assembled is adapted to be nested within combustion chamber 14. Or alternatively, central portion 54 may be bonded directly to the motor case, thus eliminating sleeve 56.

Burning of segments 38 starts on the surface 42, progresses in the L direction, and terminates at the bonded portions 54. The gap 58 between sleeve 56 and the unbonded portions of surface 44 allows for the relief of thermal stresses during combustion, and the inhibitor 50 applied on surface 44 assures that the flames do not reach the propellant until substantially all of the segrnent is burned in the L direction.

It has been determined from the combustion time versus chamber pressure curve made with this segmented grain that there are regularly occurring undulations in the pressure, apparently due to the simultaneous combustion through the overlying cell walls in successive layers of honeycomb cells. To smooth out these undulations, the individual segments may be modified in the manner shown in FIGURE 6. Here, the individual segments are made up of segment pieces 62, 64, 66, each of whose cell walls are staggered with respect to the walls of the adjacent piece. This can be done by providing a tab 72 on each segment piece and drilling a hole 74 through the tab after aligning the pieces in staggered relationship. For final assembly of the pieces, a rod (not shown) can be inserted in the hole and then the pieces can be bonded together by stringing them on the rod. Tab 72 is then cut off and the composite segment 60 can then be filled with propellant and treated in the same manner as previously described for segment 38.

The previously described internal burning propellant grains burn regressively; that is, the combustion chamber pressure decreases as the time of burning increases. This type of burning is due to the fact that burning fronts become progressively smaller and thus less gas is produced with increasing time.

In contrast to the above internal burning grains, segmented grain 76 shown in FIGURE 7 provides a grain having a substantially neutral burning curve. Segment 78 is made in the same manner as segment 38 in FIGURE 5 and has substantially the same characteristics as segment 38. Segment 80 is made by expanding the honeycomb core 82 fanwise so as to describe a semi-cylindrical surface 84 and a slightly concave surface 86 with a central depression, or concave portion 88. After filling core 82 will propellant, surfaces 84 and 86 are coated with a suitable burning inhibitor 90 and surface 88 is coated with a suitable igniter composition 92. Segment 80 is then bonded along semi-cylindrical surface 84 to a sleeve 94 in opposing relationship to segment 78. Sleeve 94 is adapted to be fitted into combustion chamber 14.

Upon ignition, segment 78 burns initially on the relatively large surface 96, and the combustion front becomes progressively smaller as it advances toward the bonded area 98. On the other hand, segment 80 burns initially on the relatively small surface 88, and its combustion front becomes progressively larger. The combination of these two segments, 78 with a regressive burning curve and 80 with a progressive burning curve, results in a composite grain with a substantially neutral burning curve.

While various embodiments of the present invention have been specifically described, other variations will be apparent to those skilled in the art, and it is to be understood that the scope of the present invention is to be limited only by the appended claims.

I claim:

1. A rocket motor comprising: a combustion chamber; an end-burning solid propellant grain in said chamber; a cellular structure in the form of a honeycomb for supporting and reinforcing the grain, said structure being so disposed in the grain that the cell axes are perpendicular to the general direction of burning of said grain.

2. A rocket motor comprising: a substantially cylindrical combustion chamber; an end-burning solid propellant grain in said chamber; a cellular structure in the form of a honeycomb for supporting and reinforcing the grain, said structure being so disposed in the grain that the cell axes are substantially perpendicular to the general direction of burning of said grain and to the combustion chamber wall, said grain having an inclined, planar ignition surface so that ignition and end-burning takes place simultaneously across several rows of the honeycomb cells.

3. A rocket motor comprising: a substantially cylindrical combustion chamber; an end-burning solid propellant grain in said chamber, said grain being comprised of semi-cylindrical half-portions bonded together, each of said portions being supported and reinforced by a cellular structure in the form of a honeycomb, said structures being disposed in each portion so that the cell axes are substantially perpendicular to the general direction of burning of said grain and said structures in each portion being disposed with relation to each other so that the respective cells are offset.

4. An end-burning solid propellant grain for a rocket motor comprising: a cylindrical metal honeycomb core having its cell axes perpendicular to the general direction of burning, said core being constituted of semi-cylindrical half-portions bonded together in such manner that the cells in one half-portion are offset with respect to the cells in the other half-portion; a solid propellant contained within said core to comprise said grain; a burning inhibitor on the cylindrical and forward surfaces of said grain; and an igniter composition on the aft surface of said grain.

5. A rocket motor comprising: a substantially cylindrical combustion chamber; a sleeve member nesting within said chamber; a plurality of internal burning propellant grain segments disposed longitudinally in said chamber; each of said segments having opposing convex surfaces, the first of said surfaces being bonded along a central portion thereof to the sleeve member, the second of said convex surfaces forming a central port; and each of said segments being reinforced with honeycomb cellular structures which are expanded fan-wise from the central portion of said first surface and in such manner that the cell ribbons are disposed in the general direction of burning.

6. A rocket motor according to claim 5, wherein the second convex surfaces are coated with an ignition composition and the unbonded portions of the first convex surfaces are coated with a burning inhibitor.

7. A rocket motor according to claim 5, wherein the respective segments are formed by a series of separate sections, each section having the walls of the honeycomb cellular structure in staggered relationship to the Walls of the cellular structure in the immediately adjacent sections.

8. A rocket motor comprising: a combustion chamber;

plurality of internal burning propellant grain segments disposed longitudinally in said chamber; each of said segments having opposing convex surfaces, the first of said surfaces being bonded along a central portion thereof to the combustion chamber wall, the second of said surfaces forming a central port; and each of said segments being reinforced with honeycomb cellular structures which are expanded fan-wise from the central portion of said first surface and in such manner that the cell ribbons are disposed in the general direction of burning.

9. A rocket motor according to claim 8, wherein the second convex surfaces are coated with an ignition composition and the unbonded portions of the first convex surfaces are coated with a burning inhibitor.

10. A rocket motor according to claim 8, wherein the respective segments are formed by a series of separate sections, each section having the walls of the honeycomb cellular structure in staggered relationship to the walls of the cellular structure in the immediately adjacent sections.

11. A rocket motor comprising: a substantially cylindrical combustion chamber; a sleeve member nesting within said chamber; a segmented, internal burning solid propellant grain having first and second segments each reinforced with a honeycomb core and disposed in the sleeve member in opposing relationship; said honeycomb cores having the cell ribbons running in the general direction of burning; said first segment having a generally semi-cylindrical surface bonded to the sleeve member and an opposing generally concave surface with a central, longitudinal depression therein; said second segment having opposing first and second convex surfaces, said first convex surface having a curvature approximating the curvature of the sleeve member and being bonded to the sleeve member along a central, longitudinal portion; an igniter composition applied to the longitudinal depression of the first segment and to the second convex surface of the second segment; and a combustion inhibitor applied to the unbonded portion of the first convex surface of the second segment, to the generally concave surface of the first segment, and to the semi-cylindrical surface of the first segment prior to bonding to the sleeve member.

References Cited by the Examiner UNITED STATES PATENT-S 2,921,521 1/1960 La Haye et a1 6035.6 2,987,880 6/1961 Kimmel 60-356 3,035,950 5/1962 Long 60-39.47 3,130,672 4/1964 Webb 6039.47

CARLTON R. CROYLE, Primary Examiner.

Claims (1)

1. A ROCKET MOTOR COMPRISING: A COMBUSTION CHAMBER; AN END-BURNING SOLID PROPELLANT GRAIN IN SAID CHAMBER; A CELLULAR STRUCTURE IN THE FORM OF A HONEYCOMB FOR SUPPORTING AND REINFORCING THE GRAIN, SAID STRUCTURE BEING SO DISPOSED IN THE GRAIN THAT THE CELL AXES ARE PERPENDICULAR TO THE GENERAL DIRECTION OF BURNING OF SAID GRAIN.
US425354A 1965-01-13 1965-01-13 Honeycomb structured propellant for rocket motors Expired - Lifetime US3316718A (en)

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Cited By (19)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3446018A (en) * 1966-12-08 1969-05-27 Thiokol Chemical Corp Liner for solid propellant rocket motor
US4655417A (en) * 1984-09-28 1987-04-07 The Boeing Company Molded ejection seat having an integrated rocket motor assembly
GB2214617A (en) * 1988-01-25 1989-09-06 Thiokol Morton Inc Rocket motors and igniters
US5024160A (en) * 1986-08-18 1991-06-18 Thiokol Corporation Rapid burning propellant charge for automobile air bag inflators, rocket motors, and igniters therefor
US5367872A (en) * 1993-04-27 1994-11-29 Thiokol Corporation Method and apparatus for enhancing combustion efficiency of solid fuel hybrid rocket motors
US6647888B1 (en) * 1999-05-06 2003-11-18 Anthony Joseph Cesaroni Oxidizer package for propellant system for rockets
US20040035312A1 (en) * 2000-11-14 2004-02-26 Biserod Hans B. Pyrotechnic charge structure
US20050115439A1 (en) * 2003-12-02 2005-06-02 Abel Stephen G. Multiple pulse segmented gas generator
US20050241294A1 (en) * 2004-04-28 2005-11-03 Cesaroni Anthony J Injector system for rocket motors
US7022196B2 (en) 2001-01-10 2006-04-04 Cesaroni Technology Incorporated Propellant system for solid fuel rocket
US20100011742A1 (en) * 2008-07-17 2010-01-21 Cavalleri Robert J Rocket Motor Containing Multiple Pellet Cells
US20100307131A1 (en) * 2009-06-09 2010-12-09 Jeremy Danforth Cartridge-Loaded Rocket Motor with Castellated Grain Segments
US20110006152A1 (en) * 2009-02-23 2011-01-13 Olden Thomas A Modular Divert and Attitude Control System
US20110024165A1 (en) * 2009-07-31 2011-02-03 Raytheon Company Systems and methods for composite structures with embedded interconnects
US8667776B2 (en) 2009-02-23 2014-03-11 Raytheon Company Pellet-loaded multiple impulse rocket motor
US8826640B2 (en) 2010-11-12 2014-09-09 Raytheon Company Flight vehicles including electrically-interconnective support structures and methods for the manufacture thereof
US9453479B1 (en) * 2007-03-22 2016-09-27 Ronald D. Jones Solid fuel grain for a hybrid propulsion system of a rocket and method for manufacturing same
US10286599B2 (en) 2007-03-22 2019-05-14 Ronald D Jones Additive manufactured thermoplastic-nanocomposite aluminum hybrid rocket fuel grain and method of manufacturing same
US10309346B2 (en) 2007-03-22 2019-06-04 Ronald D Jones Persistent vortex generating high regression rate solid fuel grain for a hybrid rocket engine

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2921521A (en) * 1958-04-25 1960-01-19 Haye Frank La Gas generator assembly
US2987880A (en) * 1958-03-17 1961-06-13 Phillips Petroleum Co Insulation for rocket motors
US3035950A (en) * 1959-02-27 1962-05-22 Phillips Petroleum Co Reinforced gelled propellants
US3130672A (en) * 1959-04-07 1964-04-28 Hexcel Products Inc Rocket grain construction

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2987880A (en) * 1958-03-17 1961-06-13 Phillips Petroleum Co Insulation for rocket motors
US2921521A (en) * 1958-04-25 1960-01-19 Haye Frank La Gas generator assembly
US3035950A (en) * 1959-02-27 1962-05-22 Phillips Petroleum Co Reinforced gelled propellants
US3130672A (en) * 1959-04-07 1964-04-28 Hexcel Products Inc Rocket grain construction

Cited By (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3446018A (en) * 1966-12-08 1969-05-27 Thiokol Chemical Corp Liner for solid propellant rocket motor
US4655417A (en) * 1984-09-28 1987-04-07 The Boeing Company Molded ejection seat having an integrated rocket motor assembly
US5024160A (en) * 1986-08-18 1991-06-18 Thiokol Corporation Rapid burning propellant charge for automobile air bag inflators, rocket motors, and igniters therefor
GB2214617A (en) * 1988-01-25 1989-09-06 Thiokol Morton Inc Rocket motors and igniters
US5367872A (en) * 1993-04-27 1994-11-29 Thiokol Corporation Method and apparatus for enhancing combustion efficiency of solid fuel hybrid rocket motors
US6647888B1 (en) * 1999-05-06 2003-11-18 Anthony Joseph Cesaroni Oxidizer package for propellant system for rockets
US20040244890A1 (en) * 1999-05-06 2004-12-09 Cesaroni Anthony Joseph Oxidizer package for propellant system for rockets
US20040035312A1 (en) * 2000-11-14 2004-02-26 Biserod Hans B. Pyrotechnic charge structure
US7117797B2 (en) * 2000-11-14 2006-10-10 Nammo Raufoss As Pyrotechnic charge structure
US7022196B2 (en) 2001-01-10 2006-04-04 Cesaroni Technology Incorporated Propellant system for solid fuel rocket
US20050115439A1 (en) * 2003-12-02 2005-06-02 Abel Stephen G. Multiple pulse segmented gas generator
US20050241294A1 (en) * 2004-04-28 2005-11-03 Cesaroni Anthony J Injector system for rocket motors
US9453479B1 (en) * 2007-03-22 2016-09-27 Ronald D. Jones Solid fuel grain for a hybrid propulsion system of a rocket and method for manufacturing same
US10286599B2 (en) 2007-03-22 2019-05-14 Ronald D Jones Additive manufactured thermoplastic-nanocomposite aluminum hybrid rocket fuel grain and method of manufacturing same
US10309346B2 (en) 2007-03-22 2019-06-04 Ronald D Jones Persistent vortex generating high regression rate solid fuel grain for a hybrid rocket engine
US20100011742A1 (en) * 2008-07-17 2010-01-21 Cavalleri Robert J Rocket Motor Containing Multiple Pellet Cells
US20110006152A1 (en) * 2009-02-23 2011-01-13 Olden Thomas A Modular Divert and Attitude Control System
US8242422B2 (en) * 2009-02-23 2012-08-14 Raytheon Company Modular divert and attitude control system
US8667776B2 (en) 2009-02-23 2014-03-11 Raytheon Company Pellet-loaded multiple impulse rocket motor
US20100307131A1 (en) * 2009-06-09 2010-12-09 Jeremy Danforth Cartridge-Loaded Rocket Motor with Castellated Grain Segments
US8281568B2 (en) * 2009-06-09 2012-10-09 Raytheon Company Cartridge-loaded rocket motor with castellated grain segments
US8809689B2 (en) 2009-07-31 2014-08-19 Raytheon Company Systems and methods for composite structures with embedded interconnects
US20110024165A1 (en) * 2009-07-31 2011-02-03 Raytheon Company Systems and methods for composite structures with embedded interconnects
US8826640B2 (en) 2010-11-12 2014-09-09 Raytheon Company Flight vehicles including electrically-interconnective support structures and methods for the manufacture thereof

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