WO2014022836A2 - Universal elliptical-sliced solid grain geometry and coupled grill-feedthrough featured assembly for solid rocket motor and coaxial hybrid rocket design - Google Patents

Abstract

A solid propellant grain geometry design and "Grill-Feedthrough" (GFT) coupled hardware assembly for advanced solid rocket motor design that embraces integration between maximized performance, through properly minimized structure weight, propulsion system features, variable thrust and random-stop. Same grain design can be extended for hybrid rocket design, applied as a variant design of current flown flight systems for advanced applications such as "continuous staging," in the context of existing and future generation space launch vehicles, ballistic missiles, antimissile's solid boosters and tactical missiles. Expected higher vehicle performance benefits are consequently achievable with very little added complexity.

Description

UNIVERSAL ELLIPTICAL-SLICED SOLID GRAIN GEOMETRY AND COUPLED GRILL-FEEDTHROUGH FEATURED ASSEMBLY FOR SOLID ROCKET MOTOR

AND COAXIAL HYBRID ROCKET DESIGN

RELATED APPLICATIONS

[0001] This application claims the benefit of priority under 35 U.S.C. §119(e) to U.S. Provisional Application No. 61/678,630 filed on August 2, 2012 entitled "Universal Elliptical-Sliced Solid Grain Geometry And Coupled Grill-Feedthrough Featured Assembly For Advanced Solid Rocket Motor Minimum Inert Weight Design, Multistage Variant Into Single Stage And Coaxial Hybrid Rocket Design," the disclosure of which is hereby incorporated by reference in its entirety.

TECHNICAL FIELD

[0002] The present disclosure relates to rocket propulsion. In one aspect it relates to solid propellant rocket motors and in particular to a novel grain configuration or geometry and its associated universal "Grill-Feedthrough" hardware (referred to hereafter as "GFT") capable of providing for random thrust control, thrust termination and restart ability features.

[0003] From another aspect it generally relates to solid rocket motors and in particular to a novel universal design procedure so intended as to be capable for providing more efficient structure weight savings, for advanced single stage equivalent design from existing multistage modifiable systems for tactical, missile defense and space industry applications.

[0004] From yet another aspect the present disclosure relates to hybrid rockets, which allows for a coaxial configuration of hybrid rockets for better overall vehicle packaging of tactical systems and structure weight reduction in space launch systems.

[0005] Relatively to all previous aspects, the present disclosure is intended as a superior common propellant grain design and associated hardware for respectively a "Grilled Solid Rocket Motor" (GSRM), or simply "Grilled Rocket Motor" (GRM) and "Coaxial Assembly Grilled Hybrid Rocket" (CAGHR) or simply "Grilled Hybrid Rocket" (GHR). BACKGROUND

[0006] It is known to any rocket designer that, generally speaking, chemical rockets have overall performance limitations and are essentially energy-limited, since the amount of energy per unit mass of propellant released during combustion is limited by the fundamental chemical behavior of propellant materials.

[0007] As is known in the field of rocket engineering, the specific impulse is the prominent measure of the velocity addition performance of a propulsion system. Since the mass of the propellant is usually a large portion of the total missile or vehicle mass, it would seem desirable to have as large value of a specific impulse as possible. For chemical rockets, this conclusion holds directly. However, maximum specific impulse does not generally mean preeminent vehicle performance. Many factors come along into play before it is possible to define the "best" engine, the "best" vehicle type (if liquid or solid propellant), the "best" shape or geometry, the "best" overall assembly (vehicle architecture), the "best" materials used, and the "best" cost.

[0008] The field of rocket engineering is quite complex, where no one can claim to have the recipe for the "best" rocket design, and specially when this type of technology has reached a point in time where regularly R&D costs often matter more than performance for its own sake. The result is an available design limit that in actuality has space for, not just evolutionary improvement, but even for new solutions for performance gain, which becomes the actual and true design limit.

[0009] Solid propulsion systems provide help in providing the thrust profile required for achieving the target terminal altitude and velocity for the payload as demonstrated in many past and current designs. Examples, in current space transportation systems are strap-on boosters, or as classic (tandem) multi-stage launch systems, e.g., the newly introduced Vega ELV (European Launch Vehicle) and American Pegasus LV. Other applications are upper- stage propulsion systems, strategic missiles and tactical missiles for antimissile programs, ground defense systems, long range operations and small diameter tactical missiles for combat aircraft all represent an example of the different fields of application of solid rocket propulsion technology.

[0010] With present day launch vehicles, entirely as multistage systems or as strap-on boosters in parallel designs, the payload capability is a small fraction of its lift-off mass, wherein the propellant mass is about 90%, structure mass (inert mass) is about 10% of the total vehicle mass in the best designed systems, in which the payload mass is also included in that 10% fraction.

[0011] The velocity addition by a stage (not including losses) is given by:

= g p In [{m_p + m_s + _pay ) / (m_s + m_pay)] = g I_sp In [(m_p / m_s + m_pay)+ 1] where g is the acceleration due to gravity, 7_sp is the specific impulse, m_p is the propellant mass of the stage, m_s the structure mass of the stage.

[0012] The burnout velocity of the vehicle depends only on the exhaust velocity, V_e, of the engine, that is the product g I_sp , and how much of the vehicle is fuel. A rocket can equal its own exhaust velocity at burnout if the mass ratio is

μ= Mo / rrif = e

where Mo is the initial mass of the vehicle, rrif its final mass and e= 2.71 8. The burnout velocity will only exceed the exhaust velocity if the mass ratio μ>ε.

[0013] The performance of a rocket depends on its structure mass m_s, and accordingly is m_pay= [(m_p / _e ^{VI g kp} - \) - m_s)

the payload for the stage (for lower stages, corresponding upper stages are the payload whereas for the final stage, the satellite is the payload). Performance can be improved any time when useful structural mass is reduced or continuously dropped. Thus, for any given propellant loading, the more we try to lower the inert mass of the stage, the higher will be the velocity addition. Thus, the necessity for a lower inert mass fraction, fj„_ert ⁼ m_s I (m_p + m_s), is a must.

[0014] Conventional solid rocket boosters designs are known as fast burning systems that produce large amounts of thrust, which still require multiple stages to reach orbital velocity or to propel long-range missile systems, increasing complexity, costs and decreasing reliability, therefore making it non practical for some applications.

[0015] Examples of typical multistage solid rocket boosters are described in U.S. Pat. No. 4,956,971, issued to Smith et al. on Sep. 18, 1990 and entitled "Solid Propellant Canister Loaded Multiple Pulsed Or Staged Rocket Motor," and U.S. Pat. No. 5,070,691, issued to Smith on Dec. 10, 1991 and entitled "Solid Propellant Canister Loaded Multiple Pulsed or Staged Rocket."

[0016] Despite attempts to improve solid rocket motor performance, driven by a demand for increased payload capability and missile range, such conventional way of construction still results in high inert mass fractions. Indeed, the propellant mass fraction is never above 94% with the best mass fractions values being for the smaller sized upper stage propulsion systems between 92% and 94%, where a 93.9% maximum value is achieved by the Star 48B spherical design configuration rocket motor of Thiokol Corporation. For very large diameter vehicles, which are 3 meters or above, the propellant mass fraction is usually never higher than 90.5% for steel cases, and 92.2% for the 3m diameter Carbon-epoxy composite 1 st stage case of the Vega ELV.

[0017] Solid rocket motors once ignited cannot change predetermined thrust or flight duration and the grain combustion continues until the entire grain is consumed. Only a repeatable, programmed variation of thrust for solid propellant motors is possible where a predetermined variation of mass flow rate has been achieved by adopting a grain geometric design which allows for changes in the propellant surface burn area at different moments during the grain combustion. Under some circumstances it is also desirable to be able to shutdown the rocket motor and thus terminate its thrust at any desired time. Once the solid fuel has been ignited it will normally burn until completely consumed. A motor that can be re-used or salvageable, while still featuring a random stop operation will have utility.

[0018] An example of a typical thrust termination assembly for rocket motor is described in U.S. Pat. No. 3,803,838, issued to Morgan et al. on Apr. 16, 1974 and entitled "Apparatus And Method Of Terminating Thrust Of A Rocket Motor." U.S. Pat. No. 2,944,390, issued to Keathley et. al. on Jul. 12, 1960 and entitled "Termination Of Thrust In Solid Propellant Rockets." U.S. Pat. No. 3,122,098, issued to Gin et. al. of NASA, on Feb.25, 1964, and Pat. Pub. No.US 2008/0092521 Al of Dulligan et al.

[0019] The shutdown and restart capability is an important safety consideration because this feature allows for an abort procedure that would otherwise result in the loss of the mission vehicle. This feature can be also useful for better vehicle energy management during orbit insertion in which current multistage vehicle systems can be substituted with "Equivalent Staged " launch vehicles with huge inert mass savings.

[0020] The typical arrangement of a modern solid rocket motor for space propulsion is illustrated in Fig. 25 (Ariane 5 P230 SM), Fig. 27 (Space Shuttle ASRM) , and Fig. 28 (a typical upper stage configuration), in which the following reference numerals designate the most common components indicated: 1-the solid propellant grain; 2-the igniter assembly; 3- the forward grain fin area, which can also be seen at aft in some stage designs; 4-the binder and EPDM insulation; 5-the forward dome or closure; 6-the central cavity port or flow passage of cylindrical shape (or cylinder-cone shaped); 7-the bolted field joints; 8-the steel (or composite) case; 9-the polar boss, always located in the aft section of the motor; 10-the nozzle assembly which often includes a TVC mechanism.

SUMMARY OF DISCLOSURE

[0021] The current state of the art technology for solid rockets and hybrid propulsion still has room for better performance and overall system improvement. Current motor designs have weight problems requiring case wall thicknesses and amounts of insulation material that results in a booster too heavy and resulting in a poor performance booster system. Some improvement in overall system performance is achieved by the booster rockets which are ejected after use, however, even such staging designs still maintain heavy equipment during flight. Thus, the cost-urgent necessity to have a solid rocket booster that is limited to one single stage that is equivalent to a tandem (2-4) multistage system to reach orbit, or to propel a long range missile system at the same distance with less propellant and overall less cost. No previous solution is universal, capable of offering applications ranging from small to large diameter rocket motor systems for overall aerospace propulsion purposes.

[0022] Current solid rocket motor design in its grain and hardware technology lacks the characteristics and functionality of a hybrid rocket, thus capable of embracing hybrid systems for improved packaging and weight performance. Such an extra feature would have tactical applications where a "two-stage" flight system (booster and sustainer) can be assembled into a whole system where the "equivalent two-stage" capability is even further improved by having a solid booster and easily throttable hybrid sustainer. Moreover, the lack of a straightforward design that can combine features with a general assembly method for a minimum weight design.

[0023] To summarize, prior disclosures lack a single universal solid rocket motor design capable of achieving minimum inert weight, motor thrust control, start-stop-restart capability, single solid stage equivalent capability to substitute multistage tandem design, through "continuous staging," applicable also for strap-on boosters, a better compact packaging approach for hybrid rockets that can require smaller lengths, improved structure weight savings, the coupling of standard solid propellant grain as a booster and a grilled-hybrid assembly as a sustainer, and finally, grain on-sight and/or long distance (in-flight) video inspection.

[0024] The present disclosure provides an advanced solid rocket propulsion system designed to have an integrated and universal assembly construction that combines the simplicity and reliability of solid propulsion systems with the features that liquid propulsion systems offer, and especially facilitates flight system weight reduction without sacrificing structural strength performance, and which permits a better packaging assembly to achieve the benefits of hybrid rocket performance.

[0025] Therefore, it is an object of this disclosure to provide for a fixed elliptical-shaped solid grain with a conveniently sliceable design on demand, "N-Grill" design, for all type of necessities, that can be used to provide an universal geometrical construction of a solid rocket motor so as to obtain a thrust characteristic curve of any desired shape while, at the same time, by turning the above mentioned weight problems into a light weight propellant housing assembly design. In fact, the present disclosure proposes a cost-convenient technology which is achievable when a medium size or especially a large diameter (3 meters or above) system propellant mass fraction can reach values as high as 97% or higher, that is 3% or less inert mass fraction, fj_nert≤ 0.03, with an optimum limit being 98%,

[0026] It is another object of this disclosure to provide for a coupled Grill-Feedthrough (hereinafter GFT) hardware for solid motor and hybrid rocket design that can be used as an integral part of said universal geometrical construction of solid rocket motors and hybrid rockets, and "continuous staging" in solid rockets applications, and which allows for thrust characteristic curves of any desired shape, while simultaneously providing for a dual random- stop feature.

[0027] It is still another object of this disclosure to provide said fixed Grill-Feedthrough (GFT) hardware for solid motor design that is capable of motor shutdown capability without disabling the motor at any point in time. This object of the disclosure is specially aimed for better safety performance during a countdown operation, an important capability of avoiding disaster, compared to current strap-on boosters or first stage motors which, once ignited on the launch pad, will continue to burn until propellant exhaustion, without regard for subsequent events. The ability of this disclosure to shutdown a booster quickly and positively during ascent also adds safety performance when this feature is, for example, integrated with a launch error detection system for strategic ballistic missiles and, if necessary, permit flight interruption without vehicle destruction, thus allowing for a better recovery mode of its pay load. The disclosure also has the same operational characteristics of multistage systems or required in-orbit multiple firings for accurate satellite orbit injections.

[0028] More particularly, but without limitations thereto, another object of the present disclosure is a straightforward, general design procedure and related assembly method, for solid motors comprising a specific universal grain design and motor assembly embodiment for overall minimized inert mass fraction, random thrust control and termination, restart capability and derived "Equivalent Staging'VContinuous Staging" concept with essential advantages. The disclosure covers a a solid rocket motor comprising a cylindrical housing configured to contain at least one grain, wherein the grain cross-sectional area is an ellipse sliced in grain elements extending along the common transversal housing axis. An embodiment includes the solid rocket motor, wherein said grains are configured in a grain design forming an ellipse as a true cross-section, when said slices are one next to the other, has a major radius b equal to a selected interior diameter of said cylindrical housing and minor radius a dependent of selected propellant burn rate, and grain design is such that grain elements define distinct fractions of the total grain mass. Another embodiment includes the solid rocket motor, further comprising at least one spacer panel between two adjacent slices of grain. Another embodiment includes the solid rocket motor, further comprising a compression ring encircling said grain. Another embodiment includes the solid rocket motor, wherein said cylindrical housing further comprises dual joint sections, one for each ending, further comprising a connecting ring between two motor segments of said cylindrical housing. Another embodiment includes the solid rocket motor, wherein an equivalent staging comprises a grain mass which equals the sum of each individual stages grain masses selected for alteration design. Another embodiment includes the solid rocket motor design, further comprising water between two adjacent slices of grain. Another embodiment includes the solid rocket motor, further comprising gas between two adjacent slices of grain. Another embodiment includes the solid rocket motor, further comprising rubber strips affixed to the exterior of said housing. Another embodiment includes the solid rocket motor, further comprising a dual system formed by external, perpendicular mounted blades for stripping of said rubber strips, and a two-blade assembly. Another embodiment includes the solid rocket motor design, further comprising a quenching liquid between two adjacent slices of grain. Another embodiment includes the solid rocket motor, further comprising a disk-shaped pusher plate coupled with a telescopic cylinder. Another embodiment includes the solid rocket motor design of claim 1 , further comprising at least one single camera system. Another embodiment includes the rocket motor wherein a hybrid rocket design comprises a central-axially positioned grain-housing with solid propellant enclosed by a coaxially surrounded oxidizer tank. The disclosure also teaches a solid rocket motor associated hardware component, comprising a system of partitions of grain elements adjacent to each other, mounted transversely to a main circular manifold connected to the interior wall of said water manifold. An embodiment includes the solid rocket motor associated hardware component, further comprising a set of linear parallel grooves, engraved on at least one side surface of said partitions in contact with at least one grain. Another embodiment includes the solid rocket motor hardware component, further comprising a linear water jet cutter made out of a groove on said partition, through which a pressurized jet is emitted perpendicularly onto each grain element surface. Another embodiment includes the solid rocket motor hardware component wherein a restart capability comprises a two-tandem consumable igniter assembly coupled to a telescopic cylinder. Another embodiment includes the solid rocket motor associated hardware component, further comprising an O-ring. Another embodiment includes the hybrid rocket, further comprising at least one pair of oxidizer injectors, symmetrically and oppositely positioned to burn exposed said grain slices. Another embodiment includes the hybrid rocket, comprising an assembly of first stage solid motor as a booster configuration coupled with a hybrid configuration as a second stage sustainer. Utility

[0029] Important advantages attained through the use of a solid rocket booster designed in accordance with the present teaching include reduced requirements insofar as reduced weight and costs of extra insulation material that is, therefore, not required inside the solid propellant housing assembly.

[0030] Such advantage of the present disclosure is the fact that the propellant grain housing section does not require an internal layer of insulation in order to protect the case from exposure to the heat of combustion, the case being the combustion chamber. Furthermore, a low combustion pressure booster, as will be the case if a fast burning rate propellant is used instead of a lower one, will require lower chamber wall thickness which in turn translates into further shaved-off structural weight.

[0031] In one embodiment of the present disclosure, such technology of high propellant mass fractions can lead to trade-off studies that aim to achieve various performance and reliability targets for future solid propulsion systems selection criteria for mission requirements.

[0032] Furthermore, a solid propulsion system of this new assembly type offers economic advantages since existing types of solid motors have their limitations. In the context of space transportation systems, when used as strap on boosters or as a single stage solid system, e.g. as an upper stage, such economic advantages can be reaped from the weight gain in cargo capacity.

[0033] As an example, in the context of the Ariane 5 space launch vehicle, with the possibility of an increase of 3.3+ tons in LEO (Low Earth Orbit) that makes for the availability of a triple satellite launch, instead of two, or simply bringing the LEO cargo capacity to approximately 25 tons of the current maximum 21 tons, would roughly translate into 10M€ ÷ 25M€ of extra gain per launch, depending on launch prices.

[0034] All systems and/or methods disclosed and claimed herein can be made and executed without undue experimentation in light of the present disclosure. While the systems and methods of this invention have been described in terms of embodiments, it will be apparent to those of skill in the art that variations may be applied to the systems and/or methods and in the steps or in the sequence of steps of assembling the system described herein without departing from the concept, spirit and scope of this disclosure. More specifically, it will be apparent that certain components of the disclosed system may be substituted for the ones described herein to achieve similar results. All such substitutions and modifications apparent to those skilled in the art are deemed to be within the spirit, scope and concept of the disclosure as defined by the appended claims.

BRIEF DESCRIPTION OF DRAWINGS

[0035] FIG. 1 is an example of a cross-sectional view, for illustrative purposes only, of a composed elliptical-sliced grain, made of 23 slices, ("23 -Grill" Design, N=23) prior to housing assembly and without any spatial separation between them.

[0036] FIG. la is an example of a cross-sectional view, for illustrative purposes only, of one-half the grain slices of FIG. 1, which compose symmetry, and shows minor and major axis a and b, with example of an individual slice of ½, area, slice width w_{s j} and thickness equal to 2r_b.

[0037] FIG. 2 is an example of a perspective longitudinal view of grain of FIG. 7 or FIG. 8, the slices of which are shown, for illustrative purposes only, compacted together in vertical position, of which half are instead shown individually separated. The whole grain is in a horizontal position, like if it were processed standing on a table ("On-Table" assembling).

[0038] FIG. 3 illustrates the whole grain of FIG. 2, not to scale, without spatial separation.

[0039] FIG. 4 is an example of a cross-sectional view, for illustrative and teaching purposes only, of grain of FIG. 1 after grain-housing assembly.

[0040] FIG. 5 is an example of a "minimum-grill" cross section (2-halfs or "2-Grill") grain design for throttling purposes of fast burning rate propellants, for small diameter tactical missiles or missile defense applications.

[0041] FIG. 6 is an example illustration of a few typical spatial separation panels for grain of

FIG. 4 and FIG. 7.

[0042] FIG. 7 is an example of a cross-sectional view, for illustrative purposes only, of the composed elliptical-sliced grain of FIG. 2, after housing assembly, consisting of an alternating sequence of separation panels.

[0043] FIG. 8 is an example of the grain cross-section view of FIG. 7 for equivalent, or new design, of single stage booster elements, based on a grain/water (or other liquid) alternating assembly, ideal for a consumable propellant grain housing and maximized weight-payload performance.

[0044] FIG. 9 is a simple illustration of a grain, after casting in vertical position, the mold panels of which are extracted by traditional means or by water injection; processing ideal for single stage boosters adopting a consumable housing. [0045] FIG. 10 is an overall longitudinal perspective of a grain housing embodiment assembly and its major components for small diameter (0.4m-1.5m) strap-on boosters applications.

[0046] FIG. 10a is a prospective view of an external compression ring with gasket for grain housing.

[0047] FIG. 10b is a prospective view of a grain joint-compression ring for elongated housings.

[0048] FIG. 11 is an overall longitudinal perspective of the grain housing embodiment assembly and its major components for large diameter (e.g., >lm-3.7m) strap-on boosters applications.

[0049] FIG. 12 illustrates a top view example of a "pusher" plate design configuration for small diameter motors, tactical missiles or applications where only a water/grain interface is selected.

[0050] FIG. 13 illustrates a top view example of a "pusher" plate design configuration for a strap-on booster application.

[0051] FIG. 14 is a prospective illustration of a consumable grain housing assembly and its major components for equivalent single stage applications and the arrow denotes its direction of motion.

[0052] FIG. 14a is a detail of a rubber strip gasket for grain/water interface.

[0053] FIG. 14b illustrates a steel cutter blade detail and relative cut-off/stripped-off, gasket action.

[0054] FIG. 15 illustrates a half-cut section prospective of a standard grain housing assembly for small diameter tactical missile applications.

[0055] FIG. 15a illustrates a prospective example of a whole grain assembly viewed from same half-cut housing skin section.

[0056] FIG. 15b shows a prospective detail portion of a silicon-rubber strip glued on the interior wall housing necessary for a tight fit and sealing purposes of the water/grain interface. [0057] FIG. 16 is an overall view of a GFT design, with exterior/interior circumferential walls details, major components and included water flow system schematic diagram for combustion flame seal-off; example for grain assembly of FIG. 7 or FIG. 8.

[0058] FIG. 17 is a detail perspective view of a standard GFT metal partition for large diameter boosters with relative, located "Primary Fluid Utility Groove" and additional "Backup," "linear seal grooves" and a propellant grain "linear- water jet cutter groove," ideal for manned flights.

[0059] FIG. 17a is a detail cross section of a "Backup Fluid Utility Groove."

[0060] FIG. 17b is a detail cross section of a "linear-water jet cutter groove" and O-ring seals.

[0061] FIG. 17c is a detail cross section of a "Primary Fluid Utility Groove" and O-ring seals.

[0062] FIG. 18 is an example of a top sectional view of a GFT and the arrow denotes its water/liquid flow pattern.

[0063] FIG. 18a is a top view cross section of the GFT internal side wall/metal plate connection.

[0064] FIG. 19 is a side sectional view of the lower portion of a GFT with in-chamber introduced propellant grain and which refers, as an example, to grain cross section of FIG. 8.

[0065] FIG. 19a is a detail cross sectional view of a GFT metal partition in between two grain slices, which is ideal for medium and large diameter motors of manned applications.

[0066 ] FIG. 19b is a detail cross sectional view of a GFT internal circumferential side wall.

[0067] FIG.19C is a detail cross sectional view of a "Compress & Seal™" O-ring for GFT internal circumferential side walls, useful also for linear applications of GFT metal partitions.

[0068] FIG.19d is a detail cross sectional view of a general backup pressure/flame seal (O- ring).

[0069] FIG. 20 is an overall perspective view of a GFT design and its major components for a Coaxial Grilled Hybrid Rocket application.

[0070] FIG. 21 illustrates a coaxial assembly for grilled hybrid rocket booster application, prior and after propellant consumption, as a redesign example for Ares 1 first stage for better performance.

[0071] FIG. 22 illustrates a perspective cross section detail view of a grain/oxidizer tank coaxial hybrid assembly.

[0072] FIG. 23 is an overall perspective view of a redesigned P230 SM strap-on booster for

Ariane 5R with "grilled-housing," combustion chamber section and GFT location.

[0073] FIG. 23a shows exterior side aft view for GFT, water tanks and thrust chamber location.

[0074] FIG. 23b is an internal side aft view for grain, GFT, tanks, and chamber line boundaries.

[0075] FIG. 23c is a schematic side view of required grain motion direction as denoted by the arrow.

[0076] FIG. 24 illustrates an overall comparison side view of a current (prior art) arrangement of the P230 SM booster of Ariane 5, and respectively its exterior/interior industrial design aspect with its redesigned "equivalent" embodiment with a maximized light weight structure.

[0077] FIG. 25 illustrates new design options for an Ariane 5R compared to P230 SM.

[0078] FIG. 26 illustrates a coaxial hybrid-single stage lower composite design for an Ariane 6.

[0079] FIG. 27 is an overall perspective internal view of an ASRJVI.

[0080] FIG. 28 is an overall perspective internal view of a typical upper-stage SRM.

[0081] FIG. 29 illustrates a comparison view of the first stage motor for NASA's Ares I and a redesigned version based on a consumable grain housing assembly, prior and after burnout.

[0082] FIG. 30 illustrates Ares V strap-on booster designs including a dual solid/coaxial assembly hybrid rocket option.

[0083] FIG. 31 are illustrative examples of the "Equivalent Staging" principle teachings (Multistage Variant into Consumable Single Stage Redesign) of the present disclosure in relation with the Vega LV, its upper-stage and fairing assembly options. [0084] FIG. 32 is an example of possible design variations of the Vega LV the side view of which comprises one-half section showing possible (of many) interior simple schematic layouts.

[0085] FIG. 33 is a simple schematic of a current Pegasus XL launch vehicle; top view looking down and side view (from the top).

[0086] FIG. 34 is a simple schematic (side view) redesign example for Pegasus XL.

[0087] FIG. 35 is a schematic comparison between a prior art Minotaur V and a few, new proposed upgraded versions consisting of an equivalent-consumable single stage and a two- stage solid-liquid.

[0088] FIG. 36 illustrates a solid and hybrid prior art rocket schematics and their respectively cross section new options (GSRM and CAGHR) for an application example of a small diameter, tactical missile Air to Air AIM 54 PHOENIX.

[0089] FIG. 37 is a simple schematic illustration of a dual consumable igniter system in its two-stage (phase 1 and 2) restart working principle, useful for coupling purposes together with the Grill-Feedthrough (GFT) hardware component.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0090] The design principle in this disclosure is different from current motors, primarily because the interior complex grain design which usually required for predetermined thrust profile is no longer necessary. However, the knowledge for motor design can be used as a reference for a preliminary design analysis, for example, Humble-Henry-Larson, "Space

Propulsion Analysis and Design," McGraw-Hill, 1995; Sutton-Biblarz, "Rocket Propulsion

Elements," 7^th ed., John Wiley & Sons, Inc., 2001.

Propellants

[0091] The solid propellant in this invention is not limited to existing compositions, even if any of the compositions previously used for conventional grains may be employed, of which too slow burn rate propellants are subjected to a slightly higher cost and lower system weight-performance which, therefore, is not recommended.

[0092] High burn rate propellant compositions are also adaptable to fit full requirements of the embodiment of the invention such as, for example, the propellant composition based on NRC-4 formulation, which is described in detail in U.S. Pat. No. 6,503,350 B2, issued to Martin et al., on Jan. 7, 2003, and entitled "Variable Burn Rate Propellant."

[0093] Furthermore, if one was to take into consideration environmental impacts from the use of current formulations, a still high performance and environmentally friendly formulation propellant for large launch vehicles is described in detail in U.S. Pat. No. 5,801 ,325, issued to Wilier et al., on Sep. 1, 1998, and entitled "High Performance Large Launch Vehicle Solid Propellants."

[0094] Another available high performance propellant, and particularly for use on tactical missile applications, has been available for more than a decade and in particular the one described in detail in U.S. Pat. No. 6,086,692, issued to Hawkins et al., of Cordant Technologies, Inc., on Jul. 1 1 , 2000, and entitled "Advanced Designs For High Pressure, High Performance Solid Propellant Rocket Motors."

Propellant Grain Geometry and Design

[0095] Referring to the drawings in detail, wherein like numerals indicate like elements, a solid grilled propellant rocket booster "elliptical-sliced grain" cross section 10 is shown according to the principles of the present disclosure, as illustrated in FIG.l, is required to be manufactured in segments, or elements, or "slices" that go through the elliptical grain entire length along its longitudinal axis, as illustrated in FIG. 2, and having minor and major radius a and b, see FIG. 1A, the manufacturing dimensions of which are dictated by engine size, design constraints (e.g., system envelope), and performance considerations of the flight system or booster.

[0096] This proposed elliptical-sliced grain design 10, "N-Grill" design, of FIG. 1, fits flight vehicles which require improved weight performance and particular feature requirements, in one embodiment of the disclosure, which is either in the context of large, medium or small diameter motors, or as tandem or parallel staged assemblies and therefore should be adopted universally, as a starting point design.

[0097] The straightforward and preferred basic elliptical-sliced grain structure of FIG. 2 is a design that fits the requirements of the disclosure, due to for easy of manufacturing, good volume efficiency, better volume to exposed grain surface area, and projected lower costs. Also, a stress relief condition for the solid grain is possible because of the absence of a case- to-grain bond interface, which represents a further advantage of this disclosure. Even more, the "N-Grilled" design allows, because of the presence of interfaces for whole grain video inspection, for either after manufacturing or quality controljnspection or long distance (inflight) control for study purposes.

[0098] The propellant material composition, its geometrical configuration through number of divisions (slices) and its feeding rate (grain consumption relative to its burning surface definition) define the motor performance characteristics and, therefore, are used to tailor the ballistic performance (thrust in function of time) of the solid rocket engine.

[0099] The number of grain slices shown in FIG. 1, is a "23-Grill" grain design which is an illustration example only, to represent a solid grain cross section without any separation between the paralleled grain sections. The 23 individual grain areas that form the whole grain elliptical area are formed by eleven symmetrical sections (slices) numbered respectively as shown, that is 12a, 12b, 12c, 12d, 12e, 12f, 12g, 12h, 12i, 12j, 12k and a middle slice 12. The traced circle 8' represents the related cylinder cross-section (and relative maximum dimensions of the housing interior wall diameter) after grain-housing assembly. At this point, it can be anticipated that any grain can be already manufactured directly into the housing and with each slice conveniently dimensioned and separated based on solid rocket design requirements.

[0100] FIG. la shows the symmetric half grain section plus middle slice cross section and illustrative example of the individual slice 12d cross section. Usually, for ease of manufacturing purposes, the major radius b should coincide with the housing interior radius, r. The value of a, which is directly proportional to propellant grain burning rate η,, instead is determined principally on design requirements, and principally motor dimensions, required propellant amount and propellant density. A trade-off study would provide the best design based on said.requirements and eventual limit costs. [0101] FIG. 2 illustrates an example of a perspective longitudinal view of grain of FIG. 7 or FIG. 8, composed of twelve slices ( a "12-Grill" grain design, N=12) one half of which is shown in an exploded view, positioned symmetrically opposite to its corresponding other half. That is, in a "grilled solid rocket motor," each propellant slice is contrary positioned if it were turned around 180° in its vertical plane. A "12-Grill" design, N=12, is thus composed of six individual slices 12a, 12b, 12c, 12d, 12e, 12f, the whole stack 22' of which is shown simplistically in FIG. 3. The chosen alternating light and dark shadows of each slice cross section (instead of a whole dark shaded surface as in FIG. 1) is to emphasize the fact that such grain geometry can be manufactured also with alternating propellants having slow and high burning rates, even if such assembly is an alternative embodiment for the present invention.

[0102] Considering the propellant grain 22', its mass M_p (in Kg), is simply A_E L_G = p_p O Vgrain , With grain ⁼ AE L_g defined as the volume of the propellant grain which obviously is also given by:

grain = M_p / p_p .

[0103] Pp is the propellant density (in kg/m³) which is (should be) uniform among all slices;

[0104] L_g (in meters) is the length (or height) of the propellant grain, which is always longer than the height of its housing because a portion, already from motor manufacturing, should be inserted inside the GFT with the grain-end inserted at the head-end of the combustion chamber.

[0105] AE is the area of the "composed" elliptical cross section, with

A_E= ab (in m²), it follows that

M_p= π Pp L_g (ab), and b= M_p Ιπ p_p L_g a or also b= V_grain / π L_g a.

Here, an appropriate trade-off between a and b can be made on a case to case basis.

[0106] Any new given engine design is specific, in which the number of grain slices, N_gs, separation panels, N_p, and their overall relative dimensioning can vary on a case to case basis depending primarily on system size and propellant burning rate ι¾. The number of separation panels is always

N_p =N_gs-l,

as can be seen in FIG. 4, FIG. 5, FIG. 7, FIG. 8, or FIG. 10, etc.

[0107] Also, based on one embodiment in the case of an even number of grain slices, an "even-grill," the total propellant mass, M_p, composed of a total

N_gs=2n

slices, can be simply summed as

M_p=2∑f₌₁ m_Pj!

because one-half of the total propellant grain slices, that is

n=(½) N_gs

are symmetric opposites in this case, as illustrated in FIG. 2.

[0108] Instead, an^'odd division ("odd-grill") has:

N_gs=2n+1

slices, with the slices separation assembled such that the one with the large width is positioned centrally along the major circle segment, exp. grain of FIG. 4, and with a mass that can be indicated with m_p o. The total mass in this case is

Thus, the widest solid grain slice coincides with the internal diameter of the grain-housing assembly.

[0109] The sum

∑f₌₁ m_P;i =m_Pjl + m_P!2...+m_P!n-1 +m_P!n of each i_th slice of propellant grain mass can be written more formally in its components.

[0110] By definition the grain slice thickness is

t_s=(a/N_gs)=2r_b (the "web" distance)

where r_b is half web distance or half slice thickness as shown in FIG. la, but should also be interpreted as being simply the burning depth in cm, relative to the grain slice thickness, that matches the propellant grain burning rate r_b and which thus matches said grain slice thickness. This is because it is more simple to design the whole grain such that each grain slice, particularly the emerged or exposed available burning surface, burns uniformly and where r_b matches the propellant burning rate, r_b, value. Of course, other complex choices are possible, and a person of ordinary skill in the art can customize the grain slice thickness based on the project's requirements. Certainly, the definition and procedure presented here, applies to any grain size and propellant burning rate which gives a symmetric and even design which is useful for any flight purpose.

[0111] Knowing the slice thickness is

a= 2 r_b N_gs.

[0112] The only dimension that varies from slice to slice is the width, more particular its midpoint, w_{s i}, of each ½ slice 12, 12a, ...12k, etc., along the cross section. Total propellant mass is:

∑P=_i m_Pji= (_Pp (2r_b) L_g)*∑¾ w_s, where w_{s j} is measured from the slice cross section center, at a distance r_b from any of the two edges, W and w₂, that is one-half of its thickness: w_s = (½) (w_\+w₂), which is half of the sum of the width of sides wiand w₂.

Again, reference is made to FIG. la. This, of course offers a quick way, by hand calculation, to obtain a very close approximation of the grain mass of each individual slice. Another better way is to determine the equivalent "true" width of a slice, w_s j _ by computing the ratio between the exact area of the slice cross section and its thickness, t_s.

[0113] Implicitly, each slice cross-section is a composition of the total area usually when a grain is made of more than two slices, as illustrated in FIG. 1 or FIG. 2:

A_S;1 ("even-grill")

A_E=nab=2r_bw_S)0+2r_b∑f₌₁ w_S;1 =A_S;0 +2∑f_=l A_S;1 ("odd-grill")

which is made (composed) of individual smaller i_t areas, A_s where

[0114] w_s o is the middle slice width;

[0115] A_{s 0} is the middle slice cross section area with A_S;i representing the surface cross section area of each individual slice, with each slice having its exposed equal opposite (or symmetric) side, with a whole thickness (the web) in cm equal to double of its manufactured burning rate (cm/sec).

[0116] Because each individual grain-slice exposed side has its unique width, w_{s i} , independently from its insertion depth, and again keeping in mind its symmetric opposite side which will have same width value, one can consider (again, for reference see FIG. 1 and FIG. la) the total sliced-grain width instead:

[0117] Most importantly, each specific solid rocket motor design has its own "propellant mass per unit length" (or "grain-length constant")

K_L=M_p/L_g (kg/m or kg/cm)

given by the ratio of the total propellant grain mass, M_p, and its length, L_g. Here, KL can be a useful factor that can be utilized for programming purposes for an eventual, to be used equipment (e.g., a measuring laser) in large diameter boosters, for the determination of the instantaneous motor thrust and history, instead of considering burn surface areas that vary in time, A_b(t). Thus, a more practical "Insertion Function" or di(t) (insertion depth that varies in time) can be useful for practical-measuring applications. A grain or booster manufacturer may customize and select which numerical program to use to obtain accurate answers of specific given design problems, for example, for a mass properties study. Finally, from the previous definition for KL, the major radius b can be also calculated as follows: K_L= _Pp ( b) → b= K_L / 2 _Pp ( r_bN_gs), which can also be used in a trade-off study.

[0118] FIG. 4 illustrates an example of the cross-section grain-housing assembly of FIG. 1 in which the elliptical-sliced propellant grain is composed of a double set, been symmetrically opposite to the central and widest slice 12, of parallel slices (or also propellant wedges) 12a, 12b, 12c, 12d, 12e, 12f, 12g, 12h, 12i, 12j, 12k that are spaced apart by "separation-guide panels" (or simply "spacer panels") 14a, 14b, 14c, 14d, 14e, 14f, 14g, 14h, 14i, 14j, 14k, in order to satisfy specific ballistic and structural design requirements for an overall housing section 22. The number of grain slices shown in FIG. 4, twenty-three, (an odd number as in FIG. 1) and relative spacer panels is a further illustration example, which is typical for slow burning rate propellants, which comprises the alternating sequence of separation thin panels or water, thus assembled into a standard cylinder-cross section which is required for flight purposes, design equivalency of multistage systems (variant redesign) or new motor design for general application purposes.

[0119] FIG. 5 illustrates an example of a "2-Grill" (or "minimum-grill") cross-section grain- housing assembly design for throttling purposes of fast burning rate propellants, which can be ideal for some particular applications of small diameter tactical missiles or missile defense interceptors with extremely high lift-off accelerations. This is the simplest assembly in which the grain-housing having wall 8 has a cross-section separated by a middle panel 14 and two half similar grains 12.

[0120] FIG. 7 and FIG. 8 illustrate examples of a "12-Grill" cross-section grain-housing designs, the number of divisions of which can be useful for slow or medium-fast burning grains, for what should be a standard assembly for any size diameter motor and type of application with the cross-section of FIG. 8 being useful also for a consumable housing application. Such design examples have (common for both) a double set of grain slices or wedges 12, 12a, 12b, 12c, 12d, 12e assembled in parallel and symmetrically opposite. Also, a double set of parallel spacer panels 14a, 14b, 14c, 14d, 14e and a middle one 14, each one composed with strips 14' and alternated with voids (or empty space 16'), complete the assembly inside cylindrical housing 8 of FIG. 7. Refer to FIG. 6 for a tridimensional view. The consumable housing application lacks spacer panels 14, 14a, etc., and instead uses, in between the grain slices 12, 12a, etc., spatial voids in parallel fashion or (partially or completely) filled with water as a spatial filler and used advantageously for the GFT operation. The water is thus pored (or filled) in between the grain wedges forming a parallel fashion of symmetric fills 16a, 16b, 16c, 16d, 16e with its center at 16. For simplicity, in the various figures, it is indicated with only 16. Prior to water fill-up, the overall grain assembly can sprayed, being separated in slices, with a coat of a thin film of silicone-type (flammable) material for, only whenever required and if necessary, impermeability unless said propellant is already manufactured with such properties.

[0121] A "3-Grill" design has a propellant burning rate, ¾ twice as fast as a "6-Grill" design and three times faster than a "9-Grill" design, etc. The "N-Grill" grain design of the present disclosure can be designed to match any standard, currently used grain geometry and has the advantage of making the overall design and casting/assembly procedures more easy, reliable and time saving, which results in an economical advantage. An advantageous and gainful grain-housing assembly in horizontal position ("on-table assembly") for potential easy grain inspection and fast, low cost process can be used either for small diameter motors for tactical missiles or medium diameter rocket motors, either for strap-on boosters or single stage elements.

[0122] FIG. 9 is a schematic illustration of a propellant casting/grain-housing assembly (in this example a "5-Grill" grain) in vertical position in which the mold panels 14*, which can be coated with a thin film of Teflon to ease the extraction process, are pulled-out from housing 8 by means of vertical sliding through an upward (or whichever is better) pulling motion 1*, or by introducing water or suitable liquid 16* for substitution with "flight-panels" (the panels intended to be used for the flight). This particular processing method is ideal for assembly of any size-diameter motor and also for consumable housings. Accordingly, must be kept in mind that the starting elliptical-sliced cross section grain design is thought to be always "turned into" a circular cross section one by the introduction of said spacer panels during the casting-assembly process or if already casted as a whole for some possible applications and in which said water is instead used in between the interfaces.

Grain-Housing Assemblies

[0123] In the present invention the "Propellant Housing" substitutes the metal or composite case of the traditional solid motor design in which said motor case, as explained in the background section, functions also as the rocket motor pressure chamber. In order to satisfy the specific design requirements the overall housing assembly 22 (see FIG. 10 and FIG. 11 or also FIG. 14, FIG. 15 and FIG. 23) should be constructed to satisfy ballistic and structural requirements. The grain-housing assembly 22 is preferably of cylindrical shape. System size, weight, acceleration flight loads and specially overall friction forces and final expected overall system assembly tightness of the propellant grain 22' must all be considered carefully into its design.

[0124] As previously explained, a variable thrust capability is provided in the solid rocket motor of the present disclosure by employing a new propellant grain design such that the segmented parts define distinct portions of the total propellant mass 22', again intended geometrically as an elliptical-sliced grain along its longitudinal axis. Said grain 22' is separated by light weight "spacer panels" 14,14a, ...14k, etc., as shown in FIG. 10, that serves the multiple function as a structural reinforcement of the overall propellant-housing assembly 22, a better thermal control of same propellant and as a guide for the solid grain drive system 32 (see for example FIG. 25).

[0125] One embodiment of the propellant-housing assembly 22 is a simple sandwiched structure of which a "4-Grill" grain design example as shown in FIG. 10, FIG. 11 and FIG. 14. In this particular example, the "4-Grill" housing is thus composed internally of a double set of parallel-symmetric propellant slices, thus symmetrically opposite to each other 12a, 12b, that are spaced apart by two symmetric 14 panels and a middle or central one 14a panel, the whole sandwiched stack of which forms the grain assembly 22. A specific parallel assembly of foam (or plastic) strips 14' covered with a thin rubber (or silicone) skin (not shown in figure) can also be used as a standard separation for large diameter housing, as shown in FIG. 6.

[0126] The spacer panels can be of any suitable light weight material with good strength-to- weight ratio and which does not support its own combustion such as, for example, any combination of thermal-plastic with interior foam or balsa wood (or any other environmentally friendly material) with a coated thermal insulation type painted surface, in which the design thickness is small, on the order of mm and based on engine size and type. Another arrangement can be internal aluminum foam panel or any other light weight foam type and exterior thin aluminum sheets or special plastic. Many possible combinations are possible in which the best light weight materials available combined with their characteristically adapted features can surely allow for acceptable light weight designs. Clearly, the material selected as spacers will depend mainly on the chosen vehicle diameter, which depends on the type of missile or vehicle application (space, ballistic or antimissile booster) and eventual design constraints (example, possible existing vehicle envelopes) that should be taken into consideration on a case to case basis. Generally, foam panels should be used in large diameter motors and only thin sheeted material for smaller diameter ones.

[0127] With respect to any medium or large diameter system, manufacturing tolerances given by design requirements should be respected in the overall new assemblies. In Fig. 10 the "4-Grill" housing 25' satisfies a light weight design requirement for a medium diameter strap-on booster. Such design is structurally characterized by several "exterior compressing rings" 24 which can be evenly distanced from an aft connection section 15a (usually to the GFT) and a top-end connection section 15j/f which is standard either for a further grain section addition or joint section (j) or a frustum (f).

[0128] The exterior of said housing 25' is also composed of an outer wall (or "skin") 8 made out of any suitable light weight composite material and several, opportunely evenly distanced, rubber O-rings 28 for necessary tight fitting between said sections. The materials suitable for the grain- housing wall 8 can be a Graphite composite, Kevlar or an Al foam between two outer thin skins of Aluminum or high strength thermal plastic, etc. A standard material for the compressing ring 24 can be instead stainless steel or D6aC steel (for lower cost) no wider than 10-20 cm, depending on booster diameter or also a Kevlar "jacket" can be used. FIG. 10a shows that such compressing ring may also have an interior rubber O-ring 29 for an appropriate and better tighter fit against housing wall 8. O-ring 29 can be either appropriately glued, and here thermal conditions that form during flight should be opportunely analyzed before appropriate choice of adhesive material, or by simply applying it inside an appropriate groove, marked by the traced center line 24', which also emphasizes the fact that in this case the thin ring 24 will be characterized by an exterior "bump" that, in this case, makes space for the required groove depth and width. The advantages of using compression rings is that they allow building a grain-housing with a very thin wall or skin, thus easily compressible by said rings. To avoid the use of compressing rings, thicker walls and accurate manufacturing tolerances in order to achieve the same required friction values. At this point, it will be clear from the context that a structural weight assessment can be helpful in obtaining a construction mode also based on motor sizes and type of application to control costs.

[0129] When necessary to join together more than one grain-housing section, for longer solid boosters, a wider "joint ring" 33 should be used together with an outer ring 24 as shown in FIG. 10. FIG. 10b illustrates the outer centrally positioned ring 24 with, for example, centrally located bump/groove 24' necessary for O-ring 29, the wider and inner joint ring 33 which should have several (in the figure indicated with traced/square-doted curved lines) rubber O-rings 28 ' for same tight fit requirements and evenly (or symmetrically) spaced apart from the main ring center and edges. The material for this wider joint ring 33 (0.5-lm) can be, for a low cost system, the same as ring 24 having 2-3 mm thickness or made in Kevlar. Thus, the rings help maintain the paralleled alternate assembly of the separation panels 14,14a, ...,14k, etc., and propellant grain slices 12,12a,...,12k, etc., (referring in general) compactly and forcedly compressed, given the fact that they must be used when the housing is made out of a thin skin, but just enough to keep the Normal force magnitude developed, perpendicular to the plane of each panel/grain intersection, hence incrementally increasing the friction force between the panel and the grain. Such friction forces, properly increased by sufficient "manufactured" compression force, naturally counterbalance each individual grain- slice weight and related inertial forces due to the vehicle acceleration.

[0130] Given the appropriate surface roughness and related static-friction force, F_s, and kinetic-friction force, F^, between panel and grain surface, of which the relative static, μ₃, and kinetic, μ^, coefficients of friction must be determined experimentally, such compression (Normal) force, N_p/_g, between panels and grain slices, can be written as follows, for an overall better understanding. By definition the static friction force is:

F_s= μ₃ Np/g and Fk= N_p/_g , where clearly μ₃ > μ^ and F_s is always proportional to the manufactured normal force, N_p/g, between the assembled grain-panel interface, in Newtons (N). The condition of equilibrium between the sum of all friction forces,∑Ff and grain weight, w_g, and total acceleration loads, w_a, is:

∑F_f - (wg + w_a) = 0.

Fk is related with the downward motion of the pusher 32. For very small insertion velocities usually the static friction force, F_s, is the only one that can be considered and it will be usually

p/g = m_grain (g+a) / μ₃ (N)

where the "g" and "a" subscripts refer to respectively the gravitational acceleration (g=9.81 m/s²) and the missile or booster acceleration. The overall contact between total grain surface, internal housing wall and interface panel assembly gives the effective surface area, SE (m ). Further details of this friction analysis and overall grain-housing mechanical study per se will not be given since they are well known by applied physicists and mechanical engineers specialists in the art. Thus, a person with ordinary skill in the art can conduct a complete flight dynamics structure analysis and related tests for a complete data collection and system behavior understanding.

[0131] Larger diameter strap-on boosters can be built similarly to assembly example 25' of FIG. 10, in which a grain-housing assembly 25, as shown in FIG. 11, has all the same components as 25' (in FIG. 11 a similar "4-Grill" example design is shown) except that several reinforcement members 26 should be added for reinforcement purposes and better overall rigidity to compensate for larger eventual motor thrust and bending moments. Such linear member 26 can be connected together with compression ring 24 at point 27 or can be made out of a whole single bar in which case the shape of compression ring 24 should be slightly modified around same point 27 to accommodate said reinforcement member 26. On the other hand, a stronger and thicker wall housing 8 can eliminate such necessity. A person with ordinary skill in the art can conduct aerospace structural analysis for a detailed assessment, with particular attention to the fact that such grain-housings are not under pressure and do not require the extra reinforcement to support internal combustion pressure. Also, small diameter tactical missiles do not require a compression ring 24 for their construction.

[0132] Any grain-housing is closed from the top by a specific and geometrically designed flat "pusher plate" 34, shown in FIG. 12, in its top view design configuration which should be a standard for small diameter motors or tactical missile applications and where only a water/grain interface is preferred such as, for example, the one shown in FIG. 8. The circumferential edge 48 is formed with several paralleled linear cuts 38 that match the interior rubber guides/seals 19 shown, as an example in FIG. 15b. Thus, for water/grain interface housing assemblies, the plate surface 36 is simply uniform and without holes except the ones described. A central connector 46, which is necessary for attachment with the hydraulic (or electric) cylinder-pusher 32 (see FIG. 22c or FIG. 24), which is intended to move longitudinally along the center axis of the housing 22 or an upper composite 102 of a rocket booster (FIG. 24 or shown by the arrow in FIG. 22c), can be simply screwed into connector 46, and said connector bolted or screwed on the plate surface by a set of screws or bolts 42. Other specific vehicle or missile designs may require a different location (e.g., exterior- symmetric for some consumable housing applications) and/or different smaller sizes of hydraulic or electric cylinders for sliced solid grains to be pushed slowly into the combustion chamber. Different propellants burn velocities will require different specific "pushes," also depending always on the type of flight application (e.g., antiballistic missiles have higher initial accelerations than space launch vehicles, etc.) and overall system sizing and assembling choice.

[0133] Any grain-housing assembly that is "dry" and thus is built with a standard whole assembly of foam/rubber strips spacer-guides 14/14' as previously explained and as shown in FIG. 6 and FIG. 7, can use a pusher plate 34 having a design configuration as shown in FIG. 13, which also illustrates its top view and is suitable for a strap-on booster application. Also in this case the pusher plate has same central connector 46 which is attached by the screw or bolt assembly 42, and with an edge 48 having same parallel linear cuts 38 but with a plate surface 36 having also paralleled linear holes 40 positioned in a tandem fashion in a single row and which match respectively the whole assembly of foam/rubber strips spacer-guides 14/14' of FIG. 6. The material to manufacture pusher plate 34 can be any suitable light weight material with suitable rigid characteristics such as high strength plastic, an aluminum foam panel or simply a thin plate of aluminum, stainless steel or light weight plastic material for smaller diameter missiles, etc.

[0134] The amount of solid propellant supplied into the thrust chamber is determined by the flight program, which takes into account the rate at which the giving propellant bums, rt,, which thus provides for a programmed computer controlled solid grain insertion velocity, v_r = di/dt (which is usually constant for constant thrust profiles) which determines the required relative insertion force, F_I? exerted by the telescopic cylinder. In one embodiment, the multistage telescopic cylinder-pusher plate assembly, which is the drive system of choice for controlled pushing and specially particularly permits smooth transition from one "burn height" to the next without major variations in chamber pressure, can be manufactured with any of the several currently available light weight aerospace materials by using, for example, Kevlar, Titanium, Al-composite or Graphite, etc. in its design since, for the reasons explained above, the necessary required forces to be exerted are very small. For example, FIG. 24 illustrates a simple schematic cut view of a strap-on booster in which its frustum has a centralized telescopic cylinder and base pusher plate.

[0135] Again, it is intended that the grain weight is "controlled" by the overall grain-housing assembly "designed" friction forces. Again, as previously explained, only a specific exploratory experimental scaled setup can precisely determine the required timely dependent insertion force magnitude, Fi (¾) by taking into consideration the following facts:

1) prior to the start of any motion the condition

Fi> F_s

is a must where the grain acceleration (extremely small) is:

a_g= (F_T - F_s)/m_g > 0

and the kinetic-friction force F^ takes over during grain motion only for a relative high velocity, which in our context should be very rare.

2) In particular, the sum of all the tangential force components,∑F_t (or the previous used F_f), is the resultant force FR, with

F_R =∑F_t = Fi - F_s = Fi (t_b) - μ₃Ν_ρ/§ , hence F_R=0→ Ff = F_s

is necessary for constant grain insertion velocities (a_g=0) and therefore constant engine thrust periods.

3) It should be taken into consideration that the resultant force FR is applied to not as a whole but composed of symmetrically located masses (usually decreasing) from the booster's center axis.

[0136] The pressure load obtained in the combustion chamber 106 and its flame temperature will not in this case deform the propellant grain 22', as is usually customary in current solid rocket motor systems, been the case also contemporarily the rocket combustion chamber, thus unaffecting the structural integrity of the grain and previous assembly of case bond system. But instead deformation will be caused, in this case, by flight acceleration loads only. All of these severe operating characteristics are, therefore, accommodated by the solid propellant rocket engine/booster design of the present disclosure.

[0137] Another embodiment of the present disclosure is to manufacture the housing such that an existing multistage rocket (or a new design) can be converted into an equivalent single-stage element in order to maximize weight savings and obtain a solid rocket motor with the best possible performance. As discussed previously, the main drawback of current case burning rocket or rocket with consumable casing is the overall mechanical complexity of their assembly and lack of full feature-ability. Again, rocket motor random thrust control and extinguishment are not discussed in the prior art in this context. The present disclosure instead, permits such technology application with the difference that a complete motor features is possible. In order to use a consumable grain-housing technology with the necessary GFT hardware, the preferred grain-housing assembly in this context should be constructed as follows.

[0138] In FIG. 14 the grain assembly 22 is glued to the housing skin 8 to form a whole assembly 23 which, also in this example is made of a "4-Grill" grain design formed with a double set of symmetric grain slices 12 and 12a, having two symmetric empty interfaces 14a (or filled with liquid, 16a) and a middle one 14 (or 16). Housing 8 is also cut or directly manufactured in a number of parallel slices. The arrow in the bottom part of FIG. 14 indicates the direction of upward motion of the GFT (not shown) relative to the consumable grain- housing. On each cut is applied along its longitudinal axis a centrally positioned rubber strip 18.

[0139] A detail of the rubber strip connection is shown in FIG. 14a where a section of same rubber strip 18 and its centered cut line 18' is positioned exactly at the center of the interface 14a/16 plane in between, in this example grain slices 12 and 12a, properly glued on top of the consumable skin housing 8 with an appropriate flammable epoxy or silicone 17. Grain slices 12 and 12a, or also any other grain slice, are internally glued to the grain-housing skin 8 with an appropriate epoxy 11. When the GFT (not shown) is put into an upward motion relative to the whole assembly 23, contemporarily each rubber strip 18 is stripped-off from the housing skin 8, as shown in FIG. 14b, by a stainless steel parallel sharp blade 53 and sliced centrally at 18' by a perpendicular stainless steel sharp blade 54 and also manufactured such that the rubber strip can break-off easily in small pieces. The two-blade assembly 53/54 is slightly wider than the rubber strip 18, mounted on top of the top-end of the GFT (GFT line) motor assembly with a number of blades assemblies 53/54 that match the number of rubber strips 18. Any GFT, in this context, should be standard (maybe reusable) and manufactured such that any number of blades can be opportunely mounted and/or changed in different positions along its top circumferential mounting base (see FIG. 18) in function of any "N-Grill" grain design. The other simple and cheaper option is to manufacture on a case by case basis, in function of the motor design.

[0140] FIG. 15 shows the basic requirement to assemble a grain-housing of a "Grilled Solid Rocket Motor" (GSRM) for small diameter tactical missiles. Such basic requirement consists of a skin housing 8 (only half section shown in the figure) having along its longitudinal axis soft rubber strips 19, positioned in parallel and distanced between them with a gauge equal to or that matches a grain slice thickness. Said soft rubber strips 19 are attached to the inner wall of housing 8 with a glue or specific epoxy 17, as shown in FIG. 15b, and thus glued entirely on one side along its entire length. The example in FIG. 15 shows seven rubber strips, thus a double set of seven for a total of fourteen rubber strips, all of the same length for a grain composed of eight slices. The details of sliced grain volume 15a is shown in FIG. 15a, in which the assembled solid grain 22 has, in this example, seven separations of which two sets are symmetric and composed of spatial interface 16a, 16b, 16c (forming a total of six) and a middle one 16. Again, eight slices of grain should be inserted inside the cylindrical envelope 8, composed of a double-symmetric set of sliced grains 12, 12a, 12b, 12c. The top-end of such assembly 22 is closed with a pusher plate similar to the one shown in FIG. 12 having its specific edge 48 made out of a total of fourteen parallel cuts 38, in this case for an "8-Grill" propellant grain design, which should be attached to an electric small telescopic cylinder or piston (not shown) for motor thrust control purposes. The down-end of the propellant grain instead is inserted into a specific sized GFT (not shown) the insertion point of which, based on the specific propellant characteristic burning rate, % starts at a certain specific distance from the grain edge indicated by the example traced line 51 '.

[0141] The combustion chamber 106 is mounted in any convenient manner at the rearward end of the GFT. A nozzle may conveniently form a continuation of said chamber 106 to generate the thrust of the rocket engine by discharging the gases generated in said combustion chamber from the propellant grain burned therein. Possibly a consumable igniter should be used to provide the start of operation of the rocket motor. Also, a small igniting charge or squib hot wire type of electrically operated igniter can be used. The variations of materials to be used depends on the type of missile application and size. The forward end of the grain- housing should have the mechanical, pneumatic or electric system to allow a piston, or plunger which is positioned to be actuated by any appropriate means, always in this case depending on the size and type of booster or missile application. Further details of the combustion chamber and nozzle section of the rocket motor per se will not be given since they are known in the art and are immaterial to an understanding of the disclosure.

Grill-Feedthrough (GFT)

[0142] Mounted within the open end of the grain-housing assembly is the GFT, its design of which is begun always as a function of the burn rate, ¾ that is characteristic of the solid propellant composition used or chosen. In one embodiment, the GFT 50 is of cylindrical shape, should always be positioned in between the combustion chamber 106 and the solid propellant grain-housing 22 (see for example FIG. 23, FIG. 24 or FIG. 36), and should be constructed from a combination of modern light weight composites and high strength/high temperature resistant alloys that attenuates temperature effects. In FIG. 16 we can see that a GFT 50, as a whole, is basically a piece of hardware composed of four main sections which includes a central grill system 150, a primary fluid (or water) tank 112' formed with an outer wall 112 (aerodynamically shaped), a high pressure rectangular-shaped circular manifold (or main annular) 116 and a sub high pressure rounded ("half-circle") inlet manifold (or secondary water manifold) 117. Sections 112', 116 and 117 together form the outer section (or "GFT-ring"). Primary tank 112' also includes the volume space for the fluid supply system which should include mainly a high pressure fluid or water pump 99, a water inlet 149, a pair of fluid/water supply tubes 98 and 97 respectively to supply primary manifold 116 (thus having an inlet 116*) and the secondary one 117 with an inlet 117* and a pair of pressure control fluid/water valves 96 and 95 respectively for manifolds 116 and 117. For small diameter missile motors, and especially for tactical applications, the "half-circle" high pressure secondary water manifold 117 is not necessary and thus also water valve 95 and supply tube 97 will not be required for such alternate (tactical) assembly. This is because said secondary circular inlet 117 is necessary only when the application of a series of "linear water-jet grain cutters" 92 (see FIG. 17 and FIG. 17b) are employed as an extra safety feature. Also, in this case, the standard design for a GFT will be made out of only three main sections, and thus will be easy to manufacture. Furthermore, a gas quenching simplified GFT design offers a cost-effective option for said small diameter tactical missiles, the principle being the same and where any suitable quenching gas substitutes said quenching fluid or water. The main standard design, instead, should be intended for all other applications and where safety and flight system reliability is a primary concern.

[0143] In order to provide the desired burning area of the propellant in said combustion chamber, the central grill system 150 (see FIG. 16 and FIG. 18) is mounted across said main circular water manifold 116. It is composed of a plurality (double set) of flat plates or partitions (specifically in-grooved or slotted) 58a, 58b, 58c, 58d, 58e and a central one 58 (in this example), and are connected to and in open communication with the interior wall 105 of water manifold 116 and said dividers are positioned to extend in generally parallel relationship to each other across the GFT 50, which forms the inlet of the combustion chamber 106. The dividers or plates 58, 58a, 58b, 58c, etc., may either be formed integrally with said circular water manifold 116 or inner wall 105 (see FIG. 16 and FIG. 18a) example by welding, or may be also fastened by screws or bolts (not shown) using a series of screw-threaded holes 103 on a reinforcing flat plate mounting 105' inside circular pressure manifold 116.

[0144] It will be understood that only the numbered such partitions 58, 58a, 58b, 58e, etc., are shown only by way of example and that any convenient number of partitions may be used, depending upon the desired size of the rocket motor or "N-Grill" design. In any type of missile application or size, it will be further understood that said partitions be parallel with each other in order to match the separated plurality of grains or slices and define the plurality of grain inlet (or outlet) openings to allow incoming grain slices 12, 12a, 12f, etc., between said partitions 58, 58a, ...,58e, etc., to come out smoothly from the housing and ensure the entrance into the combustion chamber. That is, the partitions of the GFT provide primarily a guided and safe division for the grain slices allowing for a greater burning surface area than would be available by the cross-sectional area of the original elliptical shaped grain. The surface areas of said grain wedges normally burn to provide the desired hot combustion gas and thus thrust after being ignited by any convenient igniting means. The end section of inner wall 105 of the GFT, which is attached to the combustion chamber 106 (FIG.19) should have a small space (circumferential indent) 7' in between said inner wall 105 and outermost grain slices (in this example symmetric slices 12f) for the purpose of applying an insulation material 4 (EPDM) from the beginning of the inserted and exposed grain area, basically the top of the combustion chamber area, as we can see by the examples set forth in FIG. 16, FIG. 19 and FIG. 23.

[0145] FIG. 18 shows the top sectional view of the central grill (grain inlet) 150 of a GFT, as a simple schematic that shows the water flow pattern. All partitions are shown, in this example, that is middle 58 and double set 58a, ... , 58e, in which said partitions are attached to inner wall 105, for example by a welding 142 (see FIG. 18a). The whole grill formed by partitions 58, 58a, 58e, form a double set of parallel symmetric openings 52a, 52b, 52c, 52d, 52e, 52f which match each individual propellant slice cross section area and allows said grain slices 12, 12a, 12b, 12e (not shown, see sideway FIG: 19 instead) go through and into the combustion chamber to provide for a controlled feeding. Thus a precise manufacturing tolerance must be utilized to allow for a firm and precise fit of said propellant slices into the openings. The water flows contemporarily into each partition specific slots with a flow pattern that can either go, for example, from flow path 140, from edge A to B or vice-versa, flow path 140' from edge B to A.

[0146] The use, for example, of two separate main water flow pressure valves (not shown) mounted respectively on two symmetric separation walls (not shown) should divide, in this case, the main circular water inlet at points C and D (seen from the top view in FIG.18) into two-half s symmetric and opposite pressure chambers, can allow for this feature, in which anyone of the two water valves can be used as a backup in case, for any reason, the other valve stops working. This extra feature is an alternate manufacturing option that offers a supplementary and even more reliable hardware for space launch vehicles.

[0147] The inner wall 105 forms the rectangular cross-section of main circular pressurized water manifold 116 with the outer wall 134 (see FIG. 16, FIG. 18a and FIG. 19b) and having a top cross section 138 of which the grain-housing wall 8 stands on its inner edge or indent 139 of inner wall 105 for assembling to the GFT 50. Such assembling, which requires a tight fit, is achieved with a standard set of, for example, three O-rings 28 which form attachment section 15a of the grain-housing 22, as shown previously in FIG. 10 and in detail at the top- end of FIG. 19b and FIG. 18a. Such attachment should be integrated with an appropriate high quality strength-type aerospace epoxy 146. In the case of a consumable grain-housing application, said grain-housing wall 8 will slide right in for obvious reasons and in this case the GFT 50 will not be constructed with an indent 139 to be assembled and glued to the skin housing 8. Also, the inner wall 105 can be manufactured slightly thinner, the difference being the indent depth 139. The outer fiber composite or Kevlar wall 112 (or "GFT cover") of water tank 112' can be attached from the top of cross-section 138, the outer edge of which is attached to wall 134, and in particular its left over section area 137 by using, for example, screw-threaded holes 136' (see FIG. 16 and FIG. 18a). From below it can be attached in the same manner (screwed) directly onto the lower edge of wall 134 using a second set of screw- threaded holes 112* (same as 136') and by keeping a tight water fit with the use of a circumferential O-ring 130. Arrows 49 show the assembling direction of cover 112. Such non-fixed assembling is required for easy overall construction, predominantly to ease the assembling of the fluid supply system and for safety or inspection purposes. Other possible ways for assembling cover 112 may also be utilized. The example using a screw-attaching means is for illustrative purposes, and is not meant to limit the various other fastening or attaching means which would be covered by this disclosure. The example using a screw- attaching means is for illustrative purposes, and is not meant to limit the various other fastening or attaching means which would be covered by this disclosure. Different sizes of GFT's and their related motors applications can determine a possible family-standard of construction types.

[0148] FIG. 19a shows the detail of a GFT cross section metal divider, in this example any divider 58c in between any slices 12d and 12c of a "12-Grill" solid grain design shown in FIG. 19. Each divider is coated on each side with a silicone substrate 80 (FIG.'s 16c, 16b, 16d) to allow for a tight fit of each grain slice that passes through the GFT. The top surface 74 can be in contact with either water 16 or a standard separation assembly 14 (see also FIG. 19 and FIG. 17) and be manufactured with, for example, Aluminum, Titanium, Steel or Graphite (depending on motor size and use) and surface coated with a high temperature resistant paint 79 (see also FIG. 17a, FIG. 17b and FIG. 17c) in which the silicone substrate 80 is applied. Standard construction should vary based on specific applications depending on whether small diameter tactical missiles or large diameter boosters or launch vehicles for manned or unmanned space missions are being launched. The cross section example shown in FIG. 19a, or in its perspective view of FIG. 17, can be used for large diameter flight systems because it represents the option in terms of safety (e.g., quenching features). This particular embodiment may be adapted for big size Ballistic Missiles or space launch vehicle applications in which the feature of thrust termination for emergency or the necessity of restarting the booster system ("start-stop-restart" capability) is provided in the present disclosure with a rapid means of stopping the burning of the propellant.

[0149] Still referring to FIG. 17 and FIG. 19a, another embodiment for such partition is composed, on each opposite side, of a set of specific linear paralleled slots (specifically shaped grooves or "Fluid Utility Grooves") 92, 94 and 94'. Thus, each partition has a total of six linear grooves; two linear rectangular cross-section shaped grooves 94 and 94' on each side, the upper 94 one's which should usually work as a backup and a specific linear orifice 92, one for each side, slightly positioned above the main slot or "primary utility groove" 94'. A more detailed view of cross-section of FIG. 19a is shown in FIG. 17a, FIG. 17b and FIG. 17c, and specially regarding the related O-rings distribution and type. Linear paralleled smaller O-ring grooves 70 and 72 situated at specific necessary dictated distances, which are symmetrically located on both sides of the partitions, complete such flat plate basic structure. O-ring grooves 70 should employ for their fills linear elastomeric O-rings 82' which should have the double purpose of preventing fluid or water (or extinguishing gas) from going upwards towards panel 14 and thus housing 22 or an unnecessary water-film in between the grains and partitions (reducing friction) when the primary base (or aft) linear grooves 94' are filled with pressurized water, in their function of "flame plugs" (or combustion flame seals). Furthermore, O-rings 82' should be considered for their extra utility of keeping an appropriate fit between grain and divider, which is already provided by silicone substrate 80, but still in any situation where such fit is somehow slackened, or alternatively, the O-ring can prevent undesirable cocking or blockage. For such purposes then a linear expandable-backing or stretchy 84' elastic material/O-ring is used, the function of which is to continuously "push- out" towards the grain O-ring 82' while the same gets worn out. In O-ring slots 72 instead there should be inserted linear heat-resistant O-rings 86. Many types of O-rings are usually off-the-shelf technology and any new required set can be manufactured by any of the many available companies worldwide that have great experience in such art such as, for example, Dichtomatik North America.

[0150] The linear slot 92 is geometrically shaped with a small orifice that measures 0.25-0.4 mm to function as a "linear water jet cutter" as illustrated in FIG. 19a, FIG. 17, and FIG. 17b, in the sides of said partitions thereof through which a high pressure water or also gas under pressure (e.g., carbon dioxide) can be regulated, throttled to the necessary "cutting pressure" with the high pressure water jet emitted from orifice 92 in a perpendicular direction 92' directly onto the surface grain to cut off the wedge of propellant which has been pushed into the chamber and to prevent further grain combustion. This linear water jet shaped slot 92 embodiment provides thus the means to quench the flame by cutting-off the "grain slice" also slightly above any other flame-grain surface boundary that might make its way (worst case scenario) past the top edge line of groove 94', that is already way into the GFT. A commercially available water called "Super Water" with some added chemical additives can also be used for better water pump and overall water supply system performance. A GFT used with, for example, a any type of fluid or quenching gas.

[0151] Referring again to FIG. 16, FIG. 19b and also FIG.19c and FIG.19d the propellant housing assembly 22 and GFT 50 are completely sealed off from the combustion chamber 106 by using an additional set of circumferential O-rings similarly disposed on inner wall 105. The overall assembly consists of a circumferential elastomeric O-ring 82, which is similar to the 82' O-ring with the exception of its shape (not linear in this case) and possible bigger size, which fills a circumferential slot or groove 60. Also a primary heat-resistant circumferential O-ring 86 integrated into a primary circumferential groove 61. Same as for linear O-ring 82', circumferential elastomeric O-ring 82 is also integrated together with a circumferential expandable-backing or stretchy 84 elastic material/O-ring for the same "push- out" function that satisfies all previously mentioned purposes. Similarly, the indent that forms the small space 7' left over for EPDM insulation 4 is also covered with a layer of heat- resistant 88 alloy which is ring-shaped or also can be covered appropriately with a sufficient layer of the same EPDM insulation.

[0152] Such an embodiment for the GFT provides a rapid and precise method for random quenching of the combustion when it becomes necessary to shutdown the system or allow for a restart capability. This provides a safety consideration because this permits an abort procedure to occur which that would prevent loss of the mission vehicle. It is to be understood that some variations and modifications in the manner of construction of the GFT with more than two "cutter lines" (linear jets) in the same partition or with only one single linear rectangular cross-section water groove (which is applicable for simple assembling in tactical missile applications) can be used as alternative embodiments to this disclosure.

Principles of Operation

[0153] The rocket motor is started when the propellant slices are initially ignited by an appropriate igniter already inserted into the combustion chamber. Thus said propellant slices are already inserted into the chamber for a certain depth insertion di and then, during normal operation, smoothly and continuously inserted by a telescopic cylinder by a controlled command and navigation system (not shown). Any well-known Pyrotechnic type or other known consumable igniters could be used without departing from the scope of the disclosure. A "two-staged" consumable igniter 3 is ideal (for reference see stages 1 and 2 of FIG. 37) which can allow for a restart, with its first consumable section 1 inserted into the chamber 106 and then allowing the final section 2 (second stage) to be inserted with a small telescopic cylinder 32* into the chamber 106. Said consumable igniters 3 (1 and 2) are separated by a rubber breakable membrane 400 that thus finds itself inserted and broken, 400', into the combustion chamber 106. Said dual igniter system can be attached to the internal wall 105 of the GFT through a couple of guided rods 420 which can, for example, have an extension 420' bent at 90° for attachment at point 440 of said telescopic cylinder 32*. An "end-cap" 460 works as a slider in which said rods 420 are inserted and free to move back and forth by using to respective holes (not shown) when said telescopic cylinder 32* is commanded to do so, and of which said end-cap 460 is connected rigidly to the dual igniter cylindrical structure (the exterior end of igniter 2) and works as a "stop" at the exterior edge of wall 105 after second igniter 2 insertion. Accordingly, (for reference see FIG. 23 and FIG. 25) when the engine is ignited, the multiple sliced solid propellant 22' is pushed by the telescopic cylinder 32 down the propellant housing assembly and consumed into the combustion chamber 106 as the solid material transforms into hot gases which escape through the nozzle 204. The downward and smoothly computer controlled movement of the pusher-plate 34 by said telescopic cylinder 32, which lies centrally on the longitudinal axis of the grain housing assembly 22, which is also the axis of the rocket booster 140, will force the solid grain through the GFT openings. The grain commences to burn at the surface of its exposed end. The propellant grain, depending on the required thrust, can be inserted with a speed equal to or greater than the linear burning rate of the propellant, ¾ moving axially within the housing and protruding into the free burning space where combustion takes place at all exposed unrestricted surface faces thereof, usually starting at the downstream end of the propellant plug. Accordingly, the exposed burning surface of the propellant always remains external to the GFT inlet, basically the base line of the flat metal parallel members located at the top of chamber 106.

[0154] FIG. 16 shows that the GFT can use several pressure and temperature transducers, respectively 118 (P) and 119 (T) located in any convenient location, from around the chamber wall 106 to the base or sides of the partition 58 for safety recording. Readings can be transmitted to a computer (not shown) to allow pressure modulators 93' (standard for the chamber 106) and 93 (for the safety and shutdown water system) to control the water pressure valves 96 and 95 to work properly and maintain the distributed water pressures higher than the chamber pressure. For the purpose of terminating the thrust at any desired time, it has heretofore been provided the means to do so in such a system accomplished by the GFT 50 featured assembly by pumping a suitable quencher such as water or a solution of water and soap (or also any appropriate gas) from tank 112' (which thus is also provided with water contained into interface 16) for quenching the burning, preventing the produced hot gasses from passing by the GFT O-rings assembly in which a "safe pressure" (or "flame back-up pressure") is maintained slightly higher than the combustion chamber pressure of which, as a result of the pressure differential, some water is allowed to pass into the combustion chamber 106 and vaporized. The fluid or water is essential in absorbing heat thereby keeping the partition bases from overheating. Said water flows through the linear side grooves, either in its free space (the manufacturing volume of the rectangular cross-section grooves), and onto the grain at a rate sufficient to prevent the combustion at the grain surface from progressing within said parallel flat members in an upward direction, thus forming a continuous film on the grain. When it is desired to stop the operation of the grilled rocket motor, the cylinder 32 motion should be stopped, thereby stopping the further entrance of the propellant, and contemporarily a valve could be increased in its opening allowing for the quenching water solution (or quenching gas) to extinguish immediately the burn flame at the outlet primarily initial safe line 51 of the GFT, totally "flooding" (covering) the original elliptical cross section surface area by a transverse cross-section jet area of water (or high pressure gas) previously pointed out direction 92' and/or 94* for its related 94 and 94' slots, (FIG. 17a and FIG. 17b) from the fluid utility groove openings. Although this procedure would itself terminate the burning of the fuel by virtue of the fact that the combustion of the propellant depends upon the existence of a minimum back pressure in the combustion chamber which pressure is quickly reduced below this minimum by said primary extinguishment procedure. The safety system should use for a complete reading the temperature transducers 119 with its temperature modulator 119*. If the temperature rapidly rises above seal 61, water pressure should be increased. More specifically, the temperature transducers T can be used as a switch for the water system, that is to shut it down in the last seconds of final combustion to allow for the solid grain slices to burn through (inside the GFT) in between the dividers. The pusher plate 34 functions as a top seal for continuous chamber pressure by not allowing the hot gasses to escape above the top GFT line (the base of surface 138). An appropriate O-ring (not shown) at the edge base of said pusher can be used for such purpose.

[0155] For unmanned space missions or specially manned ones, it is desired to have a method for of stopping the burning of the propellant. In the present disclosure regarding the rocket motor, this cut-off of the rocket motor thrust is achieved by opening valve 95 simultaneously with valve 96 and this command allows a high pressure water (or gas) flow to linear jet-cutter grooves 92 that form high pressure water (or gas) to come out through the orifice of said linear jet-cutter grooves 92 and contemporarily increasing the water flow rate (or gas pressure) contemporarily from all dividers, striking in a normal direction the surface of the propellant grains, consequently physically cutting off any grain residual left over from the jet cutting line, blowing them back into the combustion chamber and/or out through the nozzle thereby cutting off the burning portion of said grains and stopping combustion. Accordingly, even if, for any reason, all of the burning propellant is not removed by this cutting off through high pressure jet process, it is apparent that the possible burning area is immediately and in its greater part reduced to the generated transverse cross-sectional area of the extinguishment water outlets in the parallel divisions' arrangement. As noted above, the eventual residual small burning area is insufficient to produce the necessary pressure in chamber to support any possible combustion which might tend to continue, which eventually will be immediately after extinguished by the backup (or secondary set) of generated transverse water blasts, the complete set of outlets 94 and circumferential one 90. To restart the grilled rocket motor, for a second burn, the secondary safety/shutdown fluid system valve 95 (for the jet cutter) must be closed, the grain again introduced for a start depth, di, and the secondary consumable igniter introduced at the top of the chamber by its own insertion mechanism, as previously discussed (FIG. 37) re-igniting the grains and creating the necessary pressure for a continuous self-sustained combustion also due to the continuous grain insertion. Immediately after motor ignition the pressure in the primary safety fluid valve is opportunely increased by computer command, again with a pressure slightly above chamber pressure for normal motor operation.

[0156] The mass burning rate at any time is the product of the linear burning rate multiplied by the total surface area of the propellant grain exposed within the combustion chamber external of the GFT inlet. If the rate of advance of the grain into the chamber is increased beyond the minimum rate of the propellant, which is equal to the linear burning rate of the propellant, η,, the grain tips are being consumed still at the linear burning rate, given the particular design of the invention's grain. If the grain slices are designed with another embodiment with grain slice thickness greater than r_b (t_s>2r_b) then the grain tips can be advanced faster than being consumed at the linear burning rate. Either way, with the first embodiment, the result of this is to increase the length of the free burning grain within the chamber, that is the burning surface area. The increased length results in increased burning area within the chamber and thus a greater mass burning rate is effected, in this way by keeping the combustion always under control, even when it is free burning in the internal space of the combustion chamber. As the solid propellant burns, one section at the time, the linear dimension, that is the length of the grain, L_g, and mass of the propellant grain are decreased in time. The amount of propellant mass flow rate:

is strictly proportional to the total surface burning area, A_b, which can usually vary during the time of flight. The grain length, L_g, becomes smaller with consequently having an "insertion length," l_\ rate that is only controlled in a manner which may be either pre-programmed, a function of combustion chamber pressure, or a function of command signals received by the rocket during flight, in essence function of the required flight profile and necessary required thrust at any given moment. For each individual flight program thus the burn area varies, dAt/dt, for A_b=L_w /j(t) as seen previously in the "grain geometry and design" section, accordingly to a programmed insertion velocity, V_\— d(/i)/dt. Because the slicing technique gives us the advantage to design a large burn surface area, and specially when low speed burning propellants are implemented, the insertion velocity can be designed to be low, in the order of cm/s. For example, a redesigned Ariane 5 P230 SRB, with its t_b=130s, will have in the example of a constant thrust flight profile, a constant insertion velocity Vi =15.4 cm/s (0.154 m/s) which is equivalent to a L_g= 20m long grain, 230,000 Kg mass and a 3m diameter grain.

[0157] As previously pointed out, each specific grilled solid rocket motor design has its own propellant mass per unit length, K_L=M_p/L_g, and instead of considering a burn surface area that varies in time, the more practical insertion length that varies in time ¾t) can also be useful for practical-measuring applications, given the invention's particular sliced grain design and thus the value of KL, and so intrinsically the propellant mass rate a priori. For the P230 SRB example above, rh_p=l ,770 Kg/s. A recording for i(t) can be determined by either a direct feedback system (not shown) such as a laser device for length measuring of the telescopic cylinder extension, thus equally measuring the decrease of L_g, or - AL_g, because

that is the grain length decrease in the grain-housing equals its increase (in time) that passes through the GFT, intrinsically measuring the rate of solid fuel consumption. The amount /_1;0 is simply "the starter" length (or depth), that is the initial amount that should be kept inserted into the chamber prior to engine start (A),>AE) for reasons of easy combustion ignition.

[0158] Important to add here is that, such solid rocket motor design and operation allows, in general:

1) A_b>A_E, a typical motor operation where A =A_E (/i(t) / a) for /i(t) > a .

2) A_b<A_E, a slightly more complex motor operation and an alternate embodiment of the disclosure which requires added valves for one or more partitions to control the water flow more independently and allow for an extinguishment correlated to individual grain wedges. Obviously, here A_b =AE (/i(t) / a) with the option for ¾t) < or

3) A_b=Ae a typical motor operation for the final seconds of operation in which the grain burns independently inside the GFT, in between the partitions with the normal (A_b>Ae) thrust reduced. Such operation is also possible, that is longer thrust periods where A_b=A_E and consequently, A_b=A_E (^(t) / a) → (/i(t) / a)⁼l that is /i(t) = a . Here, said partitions should be further protected with a specific layer of EPDM insulation.

[0159] Finally, the insertion length can be considered constant for non-variable thrust profiles. Here, n_b is equal to

n_h= t_b/At_b

is nothing more than the ratio between the burning time (in seconds) and the unit time, that is At_b=l s basically a unit-less "burning number." Accordingly, this is simply done, that is to define lf=L_g/n_b, and thus relate the grain length with the propellant burning time because it is useful for the grain designer to imagine the grain divided, along its longitudinal axis, in a number of sections or segments equal to the given natural burning time. It is useful for programming purposes when an onboard computer can be used together with a laser measuring system (not shown) to measure the decrease of the propellant grain length and record the data. Theoretically, for better precision purposes, during grain length definition and related measuring design hardware, the value of n_b= t_b/At_b, can be smaller if we define the unit time in fractions of a second (example, milliseconds or 10^"3s) instead of per 1 second. Any person skilled in the art, and especially any designer of related measuring instruments or hardware, may select the use of the previous unit-less time related definitions for the use of a practical measurement purpose. It is also intended that such data collection is especially useful during new motor testing and certainly useful during space launch missions.

[0160] Extra feature: As stated previously, the volume of gas generated upon burning of the grain is directly proportional to the grain burn area and its burn rate. Prior to firing, the burn rate is controlled by the grain ambient temperature which may change with changes in storage locations of the missile, which can be housed in varying weather conditions. Therefore, to assure that the proper quantity of gas will be generated in the combustion chamber at lift-off to produce the design thrust, the burn area should be varied in inverse proportion to the changes in grain temperature and burn rate so that the volume flow rate of gas generated at lift-off is always constant. A person of ordinary skill in the art can incorporate design features to provide the means to compensate for propellant temperature changes and burn rate prior to firing to provide a constant combustion chamber pressure upon ignition of the propellant.

Coaxial Assembly Hybrid Booster

[0161] The "wagon wheel" solid fuel grain configuration used in current hybrid rockets can provide a beneficial ratio of exposed surface area to cross sectional area for the solid fuel grain. However, the wagon wheel design has disadvantages. For example, due to the slow burning rate of the fuel, the fuel grain webs become very thin during the last portion of the burn and the motor has to be shut down. This undesirably results in a high residual. It has been attempted to reinforce the wagon wheel fuel grain by incorporating solid stiffening sheets in the spoke or web portions of the grain. This too has not proven satisfactory since the fuel grain tends to separate from the solid sheets during burning. The disadvantages of such a configuration can be overcome using the hybrid rocket configuration of the present disclosure, which is another embodiment of this disclosure. FIG. 22 shows a perspective view of said an ideal embodiment of a grain/oxidizer tank coaxial assembly for a grilled hybrid rocket that is configured in accordance with the present disclosure. Currently used configurations and the design of a certain number of combustion ports is not required in this context. The principles of operation for a coaxial grilled hybrid rocket booster remain basically the same as the grilled solid rocket embodiment of the present disclosure, especially because the core or its central section has the same arrangement. A first advantage is that allows for two different grains to be assembled together, one that burns the standard way, the other that needs an oxidizer for combustion. In this way it is possible to have a first stage as a solid booster and an hybrid sustainer or vice-versa, depending on the type of application, rocket size and/or desired performance. Also, sometimes it is more convenient to adopt a shorter booster envelope. The configuration of FIG.22 also allows for further weight savings since the grain-housing section is not used contemporarily as a combustion chamber.

[0162] A small or large diameter rocket booster can use such advantageous configuration in which a grain assembly 22 conserves the same central position, surrounded by an oxidizer tank 111 which is composed of an outer Graphite /Aluminum skin wall 107 and an inner one 109 which is insulated with an insulation 115. Oxidizer tank 111 has obviously its top (and bottom) annular sections which are closed out (by welding) with any desired shaped curvature 102' (an hemispherical dome is not the best in terms of weight savings) attached to a thin skin cylindrical foil 101, made in plastic material or aluminum (for example, 1/2 mm thick) or any other appropriate light weight material with enough smoothness to accept a sliding (very low velocity) grain on its surface. Accordingly (see FIG. 21) inner tank wall 109 can either have mounted a few or more series of compressing rings 28 or not, depending on the thickness of said wall 109. Spherical shaped pressure helium bottles 102 keep the oxidizer tank 111 pressurized, thus maintaining rigidity in the overall structure. Chamber 106, nozzle 110, a specific GFT 500 having divisions 58', and an injector assembly 503/5037509 (FIG. 20) is yet another embodiment of a typical coaxial-grilled hybrid booster configuration.

[0163] Although the GFT 500 design for grilled-hybrid propulsion systems has most of its hardware components in common with a GFT (50) design for solid motors (all common components are shown in FIG. 20), one single major difference exists. Obviously, the linear water jet cutter system and the backup water grooves (as explained in the GFT section) are not necessary, yet can be included on a case to case basis. The major difference consists of additional injector assembly 503/5037509 for the oxidizer. Referring still to FIG. 20, the oxidizer should enter the top chamber flame zone 515 sideways (circumferentially), directly on the exposed grain surface by said spray injectors 509 positioned at the head top level of chamber 106, which should be designed for having solely enough volume for vaporization besides the local small-scale oxidizer diffusion on said continuously exposed grain surface, the sliced grain separation of which consequently forms a sort of upside down small channel (the distance 7 in between the webs, example see FIG. 19 and FIG. 23) which obviously goes across -the whole circular cross section, as previously discussed. Furthermore, the GFT 500 design requires for its basic completion a front injector support plate 503 for the spray assembly 509 (this can be also constructed as a whole) and any means for the oxidizer injection, which may consist, for example, of a main shutdown valve 523 to stop flow from oxidizer tank 111 when necessary, a regulating flow valve 521 controlled by a pressure modulator 525, an oxidizer pump 511, an oxidizer flow tube 505 connected with its related oxidizer inlet, and accordingly a circular manifold ("half circle" inlet jacket or wall 503') which forms the other half of the oxidizer high pressure chamber for injector assembly 509. Combustion chamber L/D ratio can be designed for an optimal mixture. Accordingly, for any preliminary design, after propellants are chosen, the oxidizer-to-fuel ratio (O/F) is determined for a given propellant weight. Any new given design can be properly iterated after analyzing simulation data. The present simplest design is the introduction of sliced-grain which allows for a constant web thickness and better volumetric efficiency (Vf_uei / V_Ch) and thus minimal volume loss. On the other hand, the common multiple port (wagon wheel) configuration, has a low 60% efficiency. From FIG. 21, we can see that after the propellant grain is inserted downwards into the chamber 106 by a telescopic cylinder mounted on the booster central axis 32', and consumed, all that is left is the lower composite (GFT and chamber) and the lightweight oxidizer tank 111. There is also an option to construct a GFT 500 having instead a pair of symmetrically opposite injectors assemblies, allowing for a lighter weight system. But a common flow pipe which is used permits continuous and contemporarily flow in both injector sections. Such design is not shown in the figures but is readily understood by a person of ordinary skill in the art, in this context. FIG. 20 shows a GFT 500 for hybrid rocket design, by having the standard GFT 50 design incorporated into it, which allows for hybrid systems to be built as two-stage (solid-hybrid) systems (see for example FIG. 36, option 2) for better rocket design application options. Furthermore, such GFT 500 assembly is still required for large diameter systems that require extra safety features.

Nozzle-Type Considerations

[0164] There is no difference in the method of operation of the main two forms, as a strap- on or "Equivalent Tandem Staged" booster system, of the disclosed new elliptical grain design application when assembled together with the overall system embodiment, the only difference being structural assembly, as previously described. In addition, the use of a conventional type nozzle, a plug nozzle or a nozzleless design only affects the nozzle efficiency of this advanced generation solid rocket motor and does not change the method of operation and its basic line performance.

[0165] In a conventional rocket design, the exhausting combustion gases are accelerated by expansion through nozzles. Efficient nozzles increase the specific impulse (kilograms of force generated per kilogram of propellant consumed per second) that a propellant provides to push the rocket forward. The specific impulse is often used in conventional rocket design as a measure of comparing rocket performances. In another possible application of the present disclosure, given the availability of fast burning rate propellants, a nozzleless design can be considered. A nozzleless rocket is described in U.S. Pat. No. 6,430,920 Bl, issued to Martin et al., on Aug. 13, 2002, and entitled "Nozzleless Rocket Motor." Such type of application would be expected to reduce substantially the performance of the rocket system, loss of which is at least partly restored by the shaved-off weight of the vehicle but with further cost savings due to a nozzleless structure. Of course, this type of application will be one in which a trade-off between performance loss and system cost requires careful evaluation. Furthermore, a plug nozzle is recommended, and is an alternative than the conventional bell-shape counterpart. System Concept Exploration Through Trade-Off Studies

[0166] As clarified before, one of the main objectives of the disclosure is to achieve better structure weight savings. Carrying out Trade-off studies of existing systems or for future systems, or also to explore a series of possible combinations of existing hardware and the various embodiments of the present disclosure is a useful first step.

It is common that during the preliminary concept design stage of a new vehicle for feasibility studies it is often useful to know how a small change in one stage will affect the performance of the rest of the vehicle. Payloads keep "growing" and this often forces booster design changes. "Old" boosters go often through continuous change to incorporate small improvements. Trade-off ratios allow us a quick method to assess performance changes when promised design values prove to be wrong. The present disclosure can bring enhanced vehicle design performance (e.g., close to its design limits) through structure weight savings, and thus also a change in the vehicles propellant mass when this is the design case chosen. No matter the approach chosen, only the first two out of the three Trade-off ratios of interest should be considered. For derivations of the following coefficient formulas and for their application, reference can be made in particular to Chapter 7 (§7.6) of William E. Wiesel, "Space Flight Dynamics," 2^nd ed., McGraw- Hill, 1997. Some application examples are shown in the "Industrial Applicability" section.

[0167] The first Trade-off ratio describes how the mass of the payload changes with changes in the structural mass of the kth stage,

mpayload _

dmsfc dAV = 0

When the last stage is considered, the one that brings the payload into orbit, is obviously:

3/Wpayload /<9miastStage = - 1

that is, the last stage structural mass trades kilogram for kilogram with payload mass. For example, a small change in the structural mass of a vehicle, 6m_si_; relates to its payload mass by

c dmpayload „

O/Wpayload * j- 0OT_si. The second Trade-off ratio describes the change in the payload mass with changes in propellant mass of the booster vehicle:

dmpayload ∑ ^~ Ve j (1/mo j - 1/mf/) + Vefe/mofe

dmpk dAV = 0 ∑i_{= 1} Vei (l/moi - 1/mfi)

This quantity will be positive, since adding fuel will increase the payload capacity.

[0168] If the present disclosures' teachings are properly applied into a new vehicle design, and only into the context of the solid rocket stages of the launch vehicle taken in assessment, no matter their order of application, a huge weight savings in structure can be achieved with significant delivered payload performance. By applying this method of preliminary investigation a person of ordinary skill in the art can obtain the advantages of payload gain from structural mass savings, for example, in the context of design of future vehicles. These are described in the upcoming section of the present disclosure where also the Trade-Off ratios are used by substituting actual current vehicle's numerical values in order to use the coefficients.

Multistage Variant Design by "Equivalent" Single Stage Element

[0169] With our current technology, it is necessary for radical changes in the way one designs rocket systems, either of liquid or solid systems. The performance of a single stage rocket is governed by the exhaust velocity Ve and the ratio of initial to final masses:

Because the performance of current chemical propulsion is limited, costs are still high and reliability is always a must and should always be improved, then one should assemble future vehicles by keeping these factors, these design considerations "fused" into a whole for new design strategy based on a simplicity criteria.

[0170] From a design standpoint, no single factor has as much influence on reliability as does design simplicity. Design simplicity not only provides increased inherent reliability but produces increased use-reliability as well. The over-all reliability of a system operating in series is the product of the individual reliabilities of the components that make up the whole system, that is R= Γι Γ₂Γ₃. . . Γ_η

where R is the over-all system reliability and n is the number of components in the system. When elements have the same reliability, then R= r".

[0171] The over-all system reliability can never exceed the reliability of the weakest component in the system. Reliability follows the product rule, and since the over-all system reliability is a product of the individual reliabilities, as components are added in series, the system reliability decreases. It can only be improved by increasing the reliability of each individual component. In many instances this is not possible. Instead over-all reliability can be increased by overdesign, which has several ways to be achieved, for example by introducing parallel components into the system, but in all cases it involves the allowance of a factor of safety or factor for contingencies. But the use of this concept has limitations.

[0172] The ignition of each stage in sequence at the staging interval is difficult to time properly, and, inevitably, some period occurs during this staging event when the control authority over the missile is at its worst. To reduce these events and improve the overall reliability of the missiles, the superpowers chose to trade performance for fewer stages.

The technologies to build event sequencers and the short duration, reproducibly timed squibs, exploding bridge-wires, or other stage separation shaped charges are costly. Serially staged missiles forces designers to carefully consider the control of a more dynamically complex vehicle. The stages and interstage breaks make the structure of a serially staged missile behave under some loading conditions as a series of smaller integral segments attached at points with flexible joints. This construction has natural frequencies that are different than a single, integral body, such as a one-stage missile. Any reduction in the number of components in the system will, by definition, increase the system's inherent reliability. Design simplicity and the reduction of the number of components in the system also means less weight or reduced inert mass fraction, reason why overdesign has its obvious limitations.

[0173] By always opting for the simplest system design the turnout is always a positive one where production costs are lower, inspection and testing techniques will be simpler, system performance is increased, and there can be created cases where such simplicity can also be felt in field operations, such as manufacturing reliability, less involved maintenance; hence, better over-all reliability of the total system. By taking this approach, the new vehicle concept which is derived approaches solid propellant launch vehicle tandem staging in a new radical way, which has been missing as a rocket performance feature for the last fifty years, since solid rocket boosters have been introduced in the space transportation field in 1961.

[0174] Because the performance of any launch system depends critically on the masses involved, the necessity to discard mass should be exercised at any opportunity. Consequently, if the present disclosure teachings of the various embodiment is appropriately applied in our context through a meticulous new assembly design among all solid stages that matches a preliminary performance analysis, indisputably the current payload capacity (estimated at 1 ,500 Kg in Polar Orbit) can be suitably increased up to a maximum design limit value which can be further investigated. Accordingly, because the amount of such research investigation is above the scopes of description of the present disclosure with its numerous possible fields of application, also because of its required necessary sophistication and length, here only a "Big Picture" and a list of possible ways to proceed in such investigation will be given, together with an indicative payload gain performance estimate, because it is useful to know how a change in a particular modification of two or three stages into one equivalent will affect the performance of the final design and its payload capacity. Given the previous reasoning, the aim for such staging application possibility is simple: to eliminate as much as possible staging all together, which is now achieved in different timing, because several motors are required to fire during the flight period in order to shave-off dead weight, which requires extra hardware, which obviously translates into extra weight and extra cost, lost performance and lost vehicle profits. A specific sequence of events must also occur for reliability of operation to occur properly.

[0175] The use of the phrases "Equivalent Staging" or "Equivalent Solid Stage" is intended as a Single Solid Stage launch vehicle that carries an amount of propellant mass only equal to the sum of the individual propellant masses of i_th stages of an existing launch vehicle or any flight system under performance alteration study by the various embodiments of the present disclosure, and in which the total burnout velocity relatively to the solid stages under examination is not altered. Therefore, under our definition, the following constraints must apply:

P_jEq =∑₌₁ m_P;1 = m_{Pj l}+ m_P;2 +... m_P;n

and

V_s,Eq ⁼∑f=_l AV_S;i=V_{s exst}

where:

M_p Eq=total equivalent solid propellant mass;

V_s Eq=total velocity of equivalent solid stage;

AV_S i = i-th velocity of existing solid stages;

V_s exst⁼ total velocity of existing solid stages.

[0176] The constraint of keeping the new designed equivalent system to have the total burnout velocity of the interested stages unchanged is because, not only such velocity is already dictated by previous mission design characteristics, and this given velocity is to be achieved, but also by the fact that here we are mainly interested in how the final payload mass changes with such fixed payload velocity. Since the payload velocity should not be allowed to change, one is interested in how the mass of the payload changes in the booster design for a fixed payload velocity. This, of course, should not be considered always the case and a vehicle designer should be able to investigate freely all performance and mission design possibility changes in function of such desired mission requirements. But, again, for the specific purposes of solely explaining the fundamental advantage of payload weight gain, which clearly translates into economic advantages, the present case study fills such requirement.

[0177] Based on a multistage vehicle's existing design data and what previously explained, it is understood that an "Equivalent Solid Stage Booster" should be designed such that its total amount of structural mass (including nozzle) is smaller than the sum of the structural masses of the stages taken into account for substitution. If the teachings in the present disclosure are properly applied into a new vehicle design, and only into the context of the solid rocket stages of the launch vehicle taken in assessment, no matter their order of application, a huge weight savings in structure- can be achieved with significant delivered payload performance.

ADDITIONAL UTILITY

Space Launch Vehicles

[0178] The past and present conventional architectural design proposals and assembly methods used to build space launch vehicles induces an unnecessary increase of their internal volume (vehicle size) and structural weight. Historically, all vehicle designers were highly motivated to decrease empty mass fraction as much as possible. Several studies at the theoretical level have been done during the fifties and sixties, at the dawn of the space age, regarding classical rocket performance through various ways of optimization practices of multistage rockets. Also, cost optimization studies are always considered. This practice also has been extended to RLV design.

[0179] There is a natural limit in the possible technological progress of chemical propulsion systems for space transportation vehicles. The main goal in studying and developing future systems should be in showing this limit. This is something that has never been achieved before and that, if ever achieved, will radically alter the way space is utilized and explored. Unfortunately, the state of the art is, theoretically, still far from realizing the full potential of spaceflight. Economics through breakthrough technology is the fundamental and major driver in how new launch vehicles should be designed and built as a practical system. The following examples chosen for applicability of the present disclosure are to vehicles currently in service or in proximity to be operative, for a better understanding of the present disclosure together with its teachings, even if such technologies, or examples, can be applied (generalized) to any type of space launcher.

Example 1: Ariane 5 P230 SRM Variant Design

[0180] Possibilities of upgrading the performance of the current flown Ariane 5 of ESA to implement its future commercial and space exploration programs can be made gradually, in order to keep the development costs to a minimum. Such upgrades should consist of keeping the same launch vehicle configuration of Ariane 5, with a vehicle architecture in which the lower composite cryogenic main stage is, in a first upgrading development phase, unchanged. The use of two ASRB's (Advanced SRB's) completes the lower composite. An improved vehicle configuration, by using ASRB's as a replacement for current used solid boosters is, for example, called "Ariane-5R" (FIG.25) which can give temporarily, sufficient vehicle transformation with safety features, which should be a prerequisite in our context towards future solid rockets as space propulsion systems, unless maximized payload fraction, that is performance for its own sake, is the only final goal. The implementation of a pair of advance solid rocket boosters will give not only a higher vehicle payload fraction (a third satellite instead of a dual launch) but also the necessary safety features which are currently unavailable for the adopted solid rocket motor technology.

[0181] There are several possible assembly options for the P230 SRM, of which all follow the teaching principles (see FIG. 25) of the present disclosure. One exemplary embodiment, in this case the less expensive to manufacture but with the least weight performance, is described here briefly and refers to assembly 120. Constraints are chosen for equivalency aspects only, that are booster main exterior dimensions, attachment points, and principle shape. Referring to FIG. 23, the combustion chamber 106 and the nozzle 110 standard assembling technology can remain unchanged of which attachment with the truncated section (aft skirt) 108 forms the lower composite 104 of the grilled solid rocket booster 120. The upper booster composite 102 and lower composite 104 are together secured in place through the GFT 50 as explained previously and furthermore by fixedly securing these elements into a whole assembly through the thrust ring 114, for example by welding, mounted on the base of the GFT 50 which in this case can have a squared annular-ring extension 50' as an extra water tank extension and for better overall assembling purposes only in the case that it is desired to have the combustion chamber 106 to have the same diameter of the GFT; otherwise, extension 50' is not necessary. This is the case, for example, of booster design option 100. Other different assembling options are possible (see FIG. 25) and the present one shown in FIG. 23 is one exemplary option. The method of assembling chambers and related nozzles are readily known and can be practiced by a person with ordinary skill in the art.

[0182] In reference to FIG. 24, which is also an example of a light weight booster version 140, it is shown that cylinder 32 is connected to the pusher plate 34 where said cylinder 32 is fixed, connected into position in the top section of the booster frustum, which coincides with the central axis of the booster in order to have longitudinal movement. The connection can consist of two necessary parts, that is the top section of said frustum can be manufactured to accept the cylinder 32 which can have an outer threaded skin and thus firmly secured into position into threaded hole 37 and secured from the outer top with a coaxial bolt 30 (or hollow-threaded cylinder). The lighter conical shaped frustum will still have space for much smaller separation solid motors 35 and a much smaller recovery parachute 39. Moreover, a necessary modified connection 31 should be used for attachment with the launcher's lower composite core. The fluid for quenching can be a simple solution of water and soap (Sutton) or a high pressure quenching gas which can be stored as already previously explained for version 120 and valid also for design 100, the one of the three proposed here that conserves most of the original design. Assembly 120 is the one that can hold more propellant mass whereas assembly 140 can be slightly more expensive to manufacture but it is also the version that can offer the best performance in terms of a lower inert mass fraction, f_inert and also because proposes a plug nozzle 204 for better ascent phase performance. From Ariane's 5 User's Manual, some of the vehicle's booster data is listed below as follows:

Mo(Ariane5-ES) = 790,000 Kg, gross mass of the launch vehicle;

77¾ρ= 240, 120 Kg, booster propellant mass;

38,210Kg, booster inert mass;

mpayioad(LEO) = 21,000 Kg, payload mass in Low Earth Orbit (LEO); and

f_iner,= 38,210/278,330= 0.137, that is 13.7% calculated inert mass fraction (f_ineit). The booster

exhaust velocity is, Ve_b= 2,750 m/s, which can be improved with better propellants. The payload-inert mass Trade gives (again, refer to the Manual for a complete data and the text reference previously cited above) for the first stage (booster burn) a coefficient value of:

which translates into a desirable payload gain value. As one example, it is possible, by properly applying the teachings of the present disclosure, to obtain a new booster design with finert =3% or less; that is shaving off some 6msl=-63,000Kg (-31 ,500Kg from each booster) to obtain:

6m_Payioad«(dm_Payioad/dmsi)6msl= (-0.052)*(-63,000)= +3,276 Kg payload gain. Properly trading, an even higher gain is feasible.

[0183] Accordingly, by shaving-off a huge 63,000kg amount of useless structure from Ariane5-ES (and surely a difference can be traded for extra propellant instead), and feasible if the present disclosure's teachings are properly applied, specially like the light-weight design 140. A new Ariane-5R design can carry a certainly beneficially -25,000 kg into low earth orbit. It must be understood that several options can be considered depending on the mission objectives, since every launcher can be assembled accordingly based on such mission requirements or final payload AV. If, for example, more propellant is required instead of more cargo capacity than the extra initial gained payload mass can be reallocated for increased propellant mass instead, either on the boosters or the core stage. The second Tradeoff ratio, that is the trade between payload and propellant, as is known in the art, can help the designer figure out a tailored option. FIG. 26 shows an example of what also an advanced "Ariane 5R" (or "Ariane 6") can look like if a set of light-weight SRB's 140 and central liquid core is varied into an "equivalent" design, a single element lower composite 160 in which the overall disclosure's teachings can be "fused" together for a coaxial-hybrid rocket booster assembly for better overall launch vehicle packaging and performance.

Example 2: Ares I First Stage and Ares V RSRM

[0184] The current available NASA state of the art for SRM technology is briefly discussed here as follows. The Ares I Crew Launch Vehicle is a two-stage vehicle which was designed principally to launch NASA's Orion CEV into low earth orbit and may also be used to launch cargo spacecraft to the ISS. The first stage of the Ares I is a five-segment reusable solid rocket motor (RSRM) derived from the four-segment boosters used in the Space Shuttle program. The added fifth segment on the Ares I solid rocket booster provides additional propellant mass and surface area to burn, providing even more thrust. This additional performance allows the launch vehicle to lift more weight, or more payload, and reach a higher altitude before the first stage separates from the upper stage, which ignites in mid- flight to propel the Orion spacecraft to Earth orbit. The addition of a fifth segment increases maximum thrust performance to approximately 3.6 million pounds, increasing total impulse by 24% over the previous existing four-segment Space Shuttle motor and enhancing vehicle and payload performance. To accommodate the additional fifth segment, certain features of the shuttle reusable solid rocket motor were modified to suit the Ares I first stage design. The motor's nozzle throat, for example, is three inches wider in diameter. The nozzle was manufactured using similar metallic materials and will perform the same functions, such as gimballing (a pivoting or swiveling mount) to move the motor nozzle, allowing the motor to point in different directions to control the vehicle's flight path. The bigger nozzle throat allows handling of the additional thrust from the five-segment booster, and meets NASA requirements to stay within the nominal operating pressure capacity of the existing steel cases.

[0185] Similar in operation to the shuttle boosters and motors, the thrust needed to lift the launch vehicle off the ground is achieved by igniting the highly-configured propellant grain, the grain geometry of which has been augmented by increasing the number of propellant fins in the forward segment from 1 1 to 12 to provide surface area to burn with a precisely controlled release of thrust, thus optimizing the thrust versus time profile. The internal propellant configuration, made of polybutadiene acrylonitrile, or PBAN, is created by pouring the propellant into an insulated and lined shuttle derived steel case segment containing grain core tooling, for molding of the grain, allowing it to solidify and then removing it. Each segment features new insulation and liner materials, incorporating the latest technology, and materials that are more environmentally friendly as well as upgraded thermal protection systems for the metal structures and seals, that provide the thermal protection required for the steel case hardware, improving overall performance.

[0186] Case segments used for the Ares I First Stage ground test have flown on a combined 48 previous Space Shuttle missions, including the aft skirt from STS-1 , bringing a rich heritage of flight-proven hardware to NASA's next-generation motor. The addition of a fifth segment brings the Ares I First Stage motor to approximately 154- 173 feet in length and is responsible for lifting the entire Ares I launch and crew vehicle stack (over two million pounds) off the ground toward Earth orbit.

[0187] From the data available on Ares 1 First Stage, either from ATK or the NASA website, the vehicle's booster data is listed below:

wibp ⁼ 648,224 Kg, booster propellant mass;

mb(inert) = 85,495 Kg, booster inert mass;

f_imr, = 0.1 165, that is 1 1,65%;

^case⁼ 58,041 Kg, case mass;

^nozzie^ 10,909 Kg, nozzle mass;

^other⁼ 16,545 Kg, miscellaneous mass.

The design is too heavy and an appropriate redesign can be applied to the booster such that the inert mass can be reduced to a minimum possible value, between 2% - 3%, especially when a burning or propulsive housing is adopted instead. An appropriate trade study can also be made in which an grilled-hybrid version can also be considered, which still retains a sufficient compact configuration and improves the overall system efficiency through a higher specific impulse (I_sp= 380s for hybrid propulsion systems). More than 60,000 Kg can be positively shaved off, where also a truncated circular aerospike nozzle should be used to complete such an advanced design, offering a better overall performance in weight savings and increased specific impulse.

[0188] It is of particular interest at this point to emphasize some facts and negative aspects regarding the conventional way of building and operation of these large size solid rocket boosters:

1 ) The optimization of thrust versus time profile of the NASA Ares 1 First Stage design, no matter how good it can be, it is still a pre-manufactured one, without any possibility of thrust variation if required.

2) The use of the motor case as a such a large (150ft long) rovent pressure vessel it is still far more dangerous than a more contained, smaller one with better possibilities for thrust control and addition of safety features due to a smaller working envelope.

3) Once ignited, there is no shutdown emergency procedure during any type of possible pre-launch inconvenience, where such feature will be desirable instead.

4) The implementation of such huge and heavy steel metal cases creates, as seen above, a remarkable loss in payload weight performance or for a better compacted vehicle design.

5) The use of the extra necessary liner and motor case insulation furthermore impacts overall weight performance.

6) Recover procedures, whenever it is necessary for flight system inspection and reuse, are complex and costly for such large size structures than for a much smaller size one; no recovery at all it is more desirable specially when the left over hardware is small and expendable.

7) Launch vehicle is way too long or tall. A shorter vehicle has less bending moment problems during flight, requires much smaller launch towers and related infrastructures and it is more adaptable for easy transportability.

[0189] Accordingly, a better vehicle packaging should be considered if a lower composite single-solid-stage, with same core diameter, is considered instead, as shown in FIG. 29 in comparison to its current design. Here, the "Equivalent-Consumable Stage" ("Continuous Staging") derived design concept 310, as the one applied for a new Vega LV design, that conserves the same nozzle and aft skirt 10, uses a combustion chamber 106 with a diameter that matches the vehicle second stage diameter with an integrated GFT motor safety and control hardware and allows better overall vehicle packaging, better payload performance and flight performance due to a reduced vehicle length and continuous length reduction due to SRM burning. In fact, after burnout what is left is only a lower composite 305 composed of the GFT, chamber 106 and aft skirt/nozzle 10, the whole structure of which can be either expended or recovered based on manufacturing specifications. The pusher and its related plate can be dropped (or not) depending on a case to case basis.

[0190] The Ares V Cargo Launch Vehicle is a two-stage, heavy-lift vehicle that NASA planned to use to carry out human missions to the Moon and other destinations. The Ares V was designed to use two five-and-a-half segment RSRMs similar to those developed for the Ares I vehicle, attached to either side of a core propulsion stage. No further trade study investigation is necessary here. The present disclosure offers essential added safety features which should be a mandatory objective for any future manned type mission. There are more than two options, besides increasing just payload or the vehicle's AV. Given the fact that this new NASA launch vehicle is already massive enough in its proportion as it is, looking similar to a Saturn V, and of which it can also match its cargo capacity, any Trade-off study should explore the potential options available:

1) Further cargo capacity increase due to weight reduction of the two RSRBs;

2) reduction of the RSRMs size, that is propellant mass at parity of existing fixed payload capacity;

3) overall increased AV, obviously useful for interplanetary missions; here, then we can trade between a fixed payload or increase propellant in the vehicle's second stage core;

4) overall reduction of vehicle size, by keeping the original sized, but redesigned, RSRBs and thus reducing the amount of core propellant; this in turn can reduce the vehicle's size, in particular its height for a better overall

packaging concept.

[0191] The principles of design and operation of a new, redesigned RSRM, will be the same as those described above, for the Ariane 5 booster, or for any strap-on booster in general as discussed previously. In our case, new designs can look very much similar to the existing one, as is shown in FIG. 30, in which the aft skirt and nozzle are conserved, for this disclosure's standard design 300 and solid/hybrid design 340. Designs 320 and dual solid/hybrid 380 are respectively the same but offer an overall better performance (with a structurally lighter and better performing plug nozzle). Payload performances are here left to be estimated on a case by case basis but with the expectation that such performance range can vary easy from 8% to 16% of any amount of shaved-off weight from its lower composite. A last consideration goes to the 46 meter (150 foot) diameter, Ares 1 drogue parachute recovery system with 900 kg (2,000 pound) mass. This larger parachute, which is deployed for safe recovery of the booster and motor components for post flight evaluation and reuse, is derived from the 41 meter (136 foot) main parachute which was used on the now retired Space Shuttle Solid Rocket Boosters and which will be used by the new five-segment solid rocket booster. Here, the benefits of having a much lighter booster structure will allow for much smaller, decreased in their weight, recovery parachutes whenever such recovery technique is used.

Example 3: VEGA ELV Single Stage Design

[0192] The previous case study made for the Ares 1 First Stage can be obviously extended as an advanced application to the Vega ELV within its specific solid propulsion 3 -stage envelope (see Prior Art in FIG. 31 and FIG. 32), or any other solid multistage vehicle.

From Vega User's Manual, the data for gross mass (g), propellant (p), inert or structure mass (s), and relative calculated propellant mass fractions (f_mp) is listed below where the subscripts 1 , 2, and 3 refers to the 1^st, 2^nd, and 3^rd stage respectively (P80FW, Z23 FW, and Z9 FW):

0.922=

92.2%

M_gj2= 25,751 Kg m_p>2 = 23,906 Kg m_Sj2 = 1 ,845 Kg f_mp>2= 0.928=

92.8%

M_gj3=l 0,948 Kg m_Pj3= 10,115 Kg m_Sj3= 833 Kg f_mP;3= 0.924= 92.4%

[0193] An "Equivalent Stage" designed to substitute the P80FW and Z23 FW (1 ^st and 2^nd stages), for example, should have a final inert mass weight much smaller than the total amount, m_s i + m_{s 2} = 9,276 Kg. The same discussion is also valid if we take into account the third Z9 FW stage, for a total of 10, 109Kg. A vehicle weight minimization is achieved by taking into consideration the following existing elements of the current Vega vehicle, which are eliminated through radical modification of the internal layout and which is followed by a new assembly, as illustrated in FIG. 31 and FIG.32 thereby reducing or eliminating the weight impact of these elements on the solid rocket booster. Between stages 1 and 2, for example, subject to an equivalency procedure, the eliminated elements will be:

1) The entire 2^nd stage structure, with nozzle and 2/3 interstage included and with it the extra cost of all this extra structure, including its stage separation system with linear cutting charge, retro rocket thrusters, electronics, etc., thus increasing the overall vehicle's reliability.

2) The majority length portion of the 1^st stage composite case with its included EPDM insulation, the 1/2 interstage section, which should be appropriately modified in shape (into cylindrical) and reduced volume, and its conventional bell nozzle which also should be opportunely substituted with a lighter and better performance given plug nozzle.

3) The 3^rd stage, separately or together, can also be included as an integrated vehicle modification, which is obviously preferred, the structure mass of which once reduced can bring into the trade-off a very close 1 to 1 weight exchange with the payload:

Smpayioad /SmiastStage = -1 and m_{s 3}= 833 Kg

→ dm_si = -833kg → δ/Wpayioad « +833kg gain.

[0194] The new vehicle concept 200' for the Vega ELV, which we can denote as "R-Vega" ("R" as "Revolutionary"), represents an advance application study, in the context of the present disclosure, of such multi-featured solid rocket motor technology, that can make consequently the new proposed Vega vehicle, an unique "modern marvel" of its kind, the rocket performance and reliability of which cannot be matched among currently used multistage vehicles.

[0195] "R-Vega" consists in one single equivalent solid component 216 with a truncated- circular aerospike nozzle 204 attached to a specific GFT 202 and connected to a central (or also an exterior set of two symmetrically positioned cylinders) telescopic cylinder 32 along the longitudinal axis of housing 22 having a consumable or propulsive skin 8. Thus the GFT/aerospike nozzle are pulled up during flight. Several other design options 200, 210, 220, and 230 are possible based on adopted faring sizes and shapes (209 and 212) the use of a solid upper stage 224 with fixed bell-shaped nozzle 226 (design 240) or a stage 222 design with an extendable nozzle 228 (design 220). If the AVUM upper stage 208 is used, it can be integrated as, for example, an interstage 208' (design 200) or without an interstage skinned structure by using a truss structure 214 (design 210).

[0196] Fig.31 shows also that such novel design is more compacted, having a reduced height (in function of the fairing size) which, for the main design 200' is of about 25m instead of the current 30.2m. The new vehicle's payload capacity (currently l ,500Kg in polar orbit) can be easily doubled as is with the equivalent-consumable 216 stage and even tripled if the AVUM stage is substituted with a better performing one, in terms of specific impulse and actually used for final total AV, instead of just mainly orbital maneuvering purposes.

Additional Application Examples: Pegasus XL and Minotaur V Space Launch Vehicles

[0197] Pegasus XL: Orbital's air-launch system represents the space industry's workhorse, providing launch services for all sorts of applications: technology demonstration, scientific investigation, remote sensing and communications missions. This small space launch vehicle was developed as an increased performance design evolution from the original Pegasus vehicle to support NASA and the USAF performance requirements, and is now the baseline configuration for all commercial launches. [0198] The launch system is a winged, three-stage, solid rocket booster (FIG. 33), with an option for a liquid fourth stage, that deploys small satellites weighing up to 1 ,000 lb (454 kg) into low-Earth orbit, weighs approximately 23,130 kg, and in a typical mission delivers its payload into orbit in a little over ten minutes. The Payload Interface Plane (PIP) is shown in the figure. Carried aloft by an L-101 1 carrier aircraft to roughly 12,000 m over open ocean, it is released in a horizontal position free-falling for five seconds before Stage 1 (SI) ignition. Stage 2 (S2) ignition occurs shortly after SI burnout, and the payload fairing is jettisoned during S2 burn as quickly as fairing dynamic pressure and payload aerodynamic heating limitations will allow, approximately 1 12,000 m and 121 seconds after vehicle airdrop. S2 burnout is followed by a long coast, during which the payload and Stage 3 (S3) achieve orbital altitude. For a non-four stage configuration, S3 then provides the additional velocity (AV) necessary to circularize the orbit, the burnout of which typically occurs about 10 minutes after launch.

[0199] Application of this disclosure's to this particular launch vehicle, for payload performance increasing purposes, the designer would simply proceed as follows:

1) Because the SI is a winged stage and from a practical/cost point of view will not make much sense modify it, unless maximum vehicle performance, quick motor shot down capabilities and full motor throttling is instead desirable, then one approach would be to apply a single stage "equivalency" design procedure between S2 and S3 (E2/3) .

2) A restart-capability should be applied since, as described above, S2 burnout is followed by a long cost of the remaining S3-payload vehicle assembly.

3) A redesigned Pegasus XL will still conserve its fourth liquid stage configuration option where, in this specific context, only a true high specific impulse engine is advised.

4) The true performance equivalency design goals should be a decrease in structure mass equivalent to the adopted S3, the payload performance gain of which simply adds to:

(3mpayload/3miastStage = - 1→ 5m_si = -123kg weight savings for S3 structure, based on the motor data from ATK's propulsion guide for the solid motor series "Orion 38." Thus

5m_Payioad » (-l)(-123kg)= +123 kg gain.

That is a straightforward 27% payload mass gain which translates into $ >1.5M gain based on current launch prices. A new redesigned Pegasus XL (exp. "Pegasus XL-R") can look schematically as shown in FIG. 34, where the "equivalent Stage 2/3" (E2/3), substitutes the original S2 and S3 stages giving a new PIP.

[0200] Minotaur V: This is a 5-stage evolutionary version (see FIG. 35) of the Minotaur IV Space Launch Vehicle (SLV) to provide cost-effective capability to launch US government sponsored small spacecraft into high energy trajectories, including Geosynchronous Transfer Orbits (GTO) as well as translunar and beyond. The first three stages of the Minotaur V are former Peacekeeper solid rocket motors. The fourth and fifth stages are commercial motors that can be selected to provide varying levels of performance. The stage four motor is a Star 48V configuration. The fifth stage can be either attitude controlled or spinning. For a spin- stabilized upper stage, a Star 37FM is used to provide maximum performance. A Star 37FMV, with gimballed, flex seal nozzle, is used for 3 -axis stabilized control. Again, trade studies can be performed to redesign a new Minotaur V (FIG. 35 is self explanatory) in which a solid-liquid two stage configuration has more potential where, giving the specific right configurations, a ^50% to 100% payload gain increase is achievable. This particular embodiment and principle of operation suggested here is similar to the "R-Vega."

Antiballistic Missiles and Tactical Missiles

[0201] The mission of the KEI (Kinetic Energy Interceptor) program is to develop and field a strategically deployable, tactically mobile, land-and sea-based capability to defeat medium- to-long-range ballistic missiles during the boost, ascent, and midcourse phases of flight. Land- and sea-mobile capabilities will use hit-to-kill technologies and a high acceleration, common booster. The initial capability addresses short - and medium - range ballistic missiles using PATRIOT Advanced Capability-3 (PAC-3) missiles, Aster 15 and Aster 30 which are of European conception that uses either a land or ship-based launching system, and also the Aegis Ballistic Missile Defense (BMD) RJM-161 Standard Missile-3 (SM-3) which is a ship- based missile system used by the US Navy. Although primarily designed as an anti-ballistic missile, the SM-3 has also been employed in an anti-satellite capacity against a satellite at the lower end of Low Earth orbit. Ground-based Interceptors (GBIs) enable engagement of intermediate-range and intercontinental ballistic missiles in the midcourse phase of their flight. The Army missile initially known as the Theater High Altitude Area Defense, or THAAD is able to engage ballistic missiles at higher altitudes and longer ranges and protect larger land areas than other terminal elements.

[0202] The above listed ABM systems represents the present state of the art and, given the fact that most of these systems use a multistage solid assembly, there is space for improvement through much better solid propulsion system assembly, and also hybrid, for multi-use capabilities for the next-generation, high performance interceptor weapon system. The present disclosure has many different applications, particularly in the quest for a better Anti Ballistic Missile defense system that allows generally for higher delivered payloads and/or increased velocities, and develop common strategies for better overall vehicle integration that leads overall better efficiency. The studies can lead to different conclusions. Design rules can turn into further numerous prospective proposals of missile design groundbreaking applications:

1) Stage Separation Tests, that are currently conducted to characterized the separation shock environment^and verify physical separation of the Stage 1 and 2 rocket motors, are no longer required for equivalency multistage applications of single solid stage systems.

2) Better overall compact packaging applications especially for larger multistage missiles and antimissiles with ground and ship based vertical launch. Air launch has also conceptually been introduced which will require thus better packaging for space savings on board the cargo-launch platform aircraft. Such application is described in detail in Pat. No. US 2007/ 0068373 Al, issued to McCantas Jr (of BAE Systems), on Mar. 29, 2007, and entitled "Air Based Vertical Launch Ballistic Missile Defense." 3) For example, an Aster 30 or SM-3 made out of a better compact packaging for an Air Based Vertical Launch in which the missile total length is significantly reduced and the missile original second stage or sustain er, remains the same, without the necessity for any changes unless there is the will to change such sustainer with an improved specific impulse small liquid engine for longer range capabilities of the missile system.

4) Better overall compact packaging for also big and smaller diameter tactical missile applications through Coaxial-Grilled Hybrid Rocket engine configurations as shown in the illustration of FIG. 22 and FIG.36 (Option 2).

5) Improved rocket performance (increased specific impulse) by adopting the Coaxial-Grilled Hybrid Rocket design.

[0203] The two basic tactical grilled motor configuration options of FIG. 36 are herein described; Option 2 uses a grilled-sliced grain configuration useful for a first higher thrust impulse (booster) shown in section SI and the remaining grain is used as the second stage S2 or sustainer integrated as an hybrid propulsion system with the use of a coaxial-oxidizer tank 111 and the appropriate GFT 500 having the two-set symmetrically opposite injectors 503 and 503'. Accordingly, a first quantity of solid grain is standard propellant while the second left over quantity is without the oxidizer component. Option 1 uses the same grain that has a first grain section SI as a booster with, for example a high burn rate propellant. S2 can have a lower burning rate propellant and thus used as a sustainer. Such grilled rocket motors and grilled hybrid systems can also be integrated with small air-breathing engines for long range and higher speed surface-to-surface and naval weapons, etc. Again, all previous teachings of the disclosure may be directly applied to form all sorts of flight systems, each one with its own different characteristics for overall better system performance. Design of new propulsion elements fused together with parallel mature technologies is the key towards innovative development that leads to an even more capable Ballistic Missile Defense System.

Strategic Ballistic Missiles

[0204] As is known in the art, ballistic missiles are missiles that have a ballistic trajectory over most of its flight path, regardless of whether or not they carry a weapon-delivery reentry vehicle and are categorized according to their range. In the last 60 years several countries have built, or sought to build, missiles with an intercontinental reach, usually under the auspices of a space launch capability. Most current long-range ballistic missiles consist of two or more stages that are stacked on top of each other and fire one at a time in sequence. Thus, extra range capability is achieved through conventional serial staging, though many effective ICBMs can be built without following any particular design guideline, and especially if financial budgets are limited and may include, besides series (or tandem) staging, parallel staging through a clustered assembling of existing liquid single-stage ballistic missiles or strap-on solid boosters.

[0205] The present disclosure will permit improvements in achievable range and/or payload performance. The current state of the art technology with respect to rocket design is "frozen" in such a way, particularly with the way that mostly serially staged missiles designs are used to deliver a payload to long distances. Examples of current "optimal," serially staged ICBMs include the U.S. Minuteman II and III, as well as the Peacekeeper missile. Each of these missiles can reach 1 1 ,000-km range and carry up to 10 nuclear warheads. To be capable of an 11,000-km range, the "ideal" ICBM would be composed of four stages, even if such design consideration is ignored, though, because of concerns about the overall reliability of the missile.

[0206] Furthermore, current American ICBMs use solid propellants. As is known from the current state of the art, the solid propellant used in their first three stages, once ignited, cannot be extinguished; it burns until exhaustion. This particular aspect and especially, for example, when strategic systems are taken into consideration given the nature of the type of weapons involved, can be positively addressed in future systems designs once an appropriate anytime emergency shutdown system is integrated into a solid stage element for any type intrinsic or caused missile launch system problem. To this end, the disclosure provides such safety propulsion device which is very useful to avoid any kind of potential uncontrollable disaster (e.g., an accidental launch error). Examples also include propulsion devices which are already on a launch silo or soon after breaking the ocean water surface after a submarine launch. The universal configuration in the present disclosure provides overall performance, safety features and reliability.

[0207] Taken into consideration all the teachings of this disclosure, an embodiment of an equivalent single-stage propulsion motor should be considered in favor for substitution of any given multistage vehicles. For a given amount of propellant, range may be sacrificed to increase payload and vice-versa. For a fixed burnout altitude, the vehicle's payload may be maximized by minimizing the burnout speed. In sharp contrast to conventional solid propellant rockets, the present disclosure provides the described efficient method for variable thrust control. This makes it possible to design new solid rocket motors which can be used to launch a payload at varying operating distances by regulating the amount of solid propellant that is consumed, and such control parameter, if combined with the variation in the launch angle, allows for manufacturing of a fewer number of tactical motor sizes to offer a wide range of performance capabilities in regards to flight trajectory and effective range. A discussion for reference purposes on "Ballistic Missile Trajectories" can be seen in particular in Bate, Roger R., Donald D. Mueller, and Jerry E. White, Fundamentals of Astrodynamics, pp. 277-320. New York: Dover Publications, 1971.

[0208] The teachings in this disclosure has applicability to solid and hybrid rockets and permit one of skill in the art to explore the technological limit of what is possible. It is to be understood that variations in the manner of operation and construction of the various embodiments of the disclosure disclosed and claimed herein can be made and executed without undue experimentation in light of the present disclosure. While the construction, application of such construction has been described in terms of embodiments, it will be apparent to those of skill in the art that variations may be applied to the methods of construction described herein without departing from the concept, spirit and scope of the disclosure. More specifically, it will be apparent that certain modifications may be substituted to achieve similar results. All such modifications apparent to those skilled in the art are deemed to be within the spirit, scope and concept of the invention as defined by the appended claims.

Claims

CLAIMS What is claimed is:

1. A solid rocket motor comprising a cylindrical housing configured to contain at least one grain, wherein the grain cross-sectional area is an ellipse sliced in grain elements extending along the common transversal housing axis.

2. The solid rocket motor of claim 1 , wherein said grains are configured in a grain design forming an ellipse as a true cross-section, when said slices are one next to the other, has a major radius b equal to a selected interior diameter of said cylindrical housing and minor radius a dependent of selected propellant burn rate, and grain design is such that grain elements define distinct fractions of the total grain mass.

3. The solid rocket motor of claim 1 , further comprising at least one spacer panel between two adjacent slices of grain.

4. The solid rocket motor of claim 1 , further comprising a compression ring encircling said grain.

5. The solid rocket motor of claim 1 , wherein said cylindrical housing further comprises dual joint sections, one for each ending, further comprising a connecting ring between two motor segments of said cylindrical housing.

6. The solid rocket motor of claim 1 , wherein an equivalent staging comprises a grain mass which equals the sum of each individual stages grain masses selected for alteration design.

7. The solid rocket motor of claim 1 , further comprising water between two adjacent slices of grain.

8. The solid rocket motor of claim 1 , further comprising gas between two adjacent slices of grain.

9. The solid rocket motor of claim 1 , further comprising rubber strips affixed to the exterior of said housing.

10. The solid rocket motor of claim 1 , further comprising a dual system formed by external, perpendicular mounted blades for stripping of said rubber strips, and a two- blade assembly.

1 1. The solid rocket motor of claim 7, further comprising a quenching liquid between two adjacent slices of grain.

12. The solid rocket motor of claim 1 , further comprising a disk-shaped pusher plate coupled with a telescopic cylinder.

13. The solid rocket motor of claim 1 , further comprising at least one single camera system.

14. The rocket motor of claim 1, wherein a hybrid rocket design comprises a central- axially positioned grain-housing with solid propellant enclosed by a coaxially surrounded oxidizer tank.

15. A solid rocket motor associated hardware component, comprising a system of partitions of grain elements adjacent to each other, mounted transversely to a main circular manifold connected to the interior wall of said water manifold.

16. The solid rocket motor associated hardware component of claim 12, further comprising a set of linear parallel grooves, engraved on at least one side surface of said partitions in contact with at least one grain.

17. The solid rocket motor hardware component of claim 14, further comprising a linear water jet cutter made out of a groove on said partition, through which a pressurized jet is emitted perpendicularly onto each grain element surface.

18. The solid rocket motor hardware component of claim 14, wherein a restart capability comprises a two-tandem consumable igniter assembly coupled to a telescopic cylinder.

19. The solid rocket motor associated hardware component of claim 14, further comprising an O-ring.

20. The rocket motor of claim 14, further comprising at least one pair of oxidizer injectors, symmetrically and oppositely positioned to burn exposed said grain slices.

21. The rocket motor of claim 14, comprising an assembly of first stage solid motor as a booster configuration coupled with a hybrid configuration as a second stage sustainer.