Background of the Invention Solid propellant rocket motors are used in a wide variety of applications. Solid propellants provide advantages over liquid fueled motors in ease of construction, simplicity of design, pre-launch safety, and performance. However, despite these advantages, there are many circumstances in which solid rocket motors have traditionally not been preferable alternatives to liquid fueled rocket motors, or combination systems of liquid and solid rockets.

One such application is the launching of payloads into orbit around the Earth. In order to place a payload in orbit, the launching system must not only lift the payload out of the Earth's atmosphere, but must accelerate it to a speed such that it will remain in motion around the Earth without travelling back into the atmosphere and falling to the surface.

Traditionally, a system designed to reach the needed speed and altitude has always required multiple rocket engines, operating in a sequential fashion. Each group of engines which operate simultaneously is referred to as a "stage". For example, NASA's Gemini program used separate stages of liquid fueled engines in order to propel a capsule into an Earth orbit. When the first group of engines (the first stage) exhausted their fuel, they were jettisoned and the next set of engines (the second stage) were ignited. The Space Shuttle uses liquid fueled main engines throughout acceleration to orbit, as well as solid fueled booster rockets for the initial phases of the launch. A similar system is used in Boeing's Delta launch vehicle, which uses nine solid fuel booster motors and a liquid fueled main engine.

In these and all previous systems, multiple stages of engines burning at different times are necessary in order to achieve the needed thrust at all points along the flight path to properly insert the payload into Earth orbit. No single engine or engine stage has been capable of providing the necessary thrust at all points along the flight path, thereby necessitating the use of multiple stages and booster rockets.

A vehicle which is capable of launching a payload into orbit using only a single stage of engines is referred to as a"single-stage-to-orbit"vehicle (SSTO). A SSTO has certain advantages over a multi-stage vehicle in matters of design, construction, and efficiency.

A multi-stage vehicle uses at least one engine in each stage. This means that a multi stage vehicle generally requires the use of more individual engines than a SSTO would. Each additional engine adds mass and complexity to the vehicle, generally making it more expensive. Furthermore, any engine which is not operating for a given part of the launch profile is dead weight that the engines which are operating must work to accelerate.

A further disadvantage of a multi-stage design is that when a stage has finished burning all of its fuel, it either must be carried on to orbit as dead weight producing no thrust, or it must be jettisoned. Jettisoning makes the design more complicated and wasteful, and makes construction and testing more difficult because systems for separating and jettisoning the expended stages must be added to the design.

For these reasons, it is advantageous to produce a single stage design which is capable of reaching orbit.

However, in order to construct a SSTO, certain qualities must be built into the engines which will be used. Among these are high thrust, light weight, variable levels of thrust, and the flexibility to operate under conditions from high ambient pressure at launch, all the way to vacuum conditions, such as are experienced in orbit. Although various engines have these various properties, it is difficult to design a single engine with all of these properties. This is why traditional orbital launch systems utilize multi-stage rocket engines, often mixing both liquid and solid fueled engines.

Solid rocket engines generally have a higher effective thrust than liquid fueled rocket engines. Although the combustion in a liquid fueled rocket is more chemically efficient, there are many additional systems that are used in a liquid fueled rocket when compared to a solid fuel rocket motor. Most of these have to do with the storage, pressurization, and pumping of the liquid fuel within the engine itself, as well as controlling the continuous burning while pumping new fuel into the combustion chamber.

By contrast, a solid fueled rocket engine is quite simple. A solid mixture of fuel and oxidizer is formed and placed into an interior cavity of a rocket motor casing. Once ignited, the solid propellant mixture burns away from its exposed surface, much as any ordinary solid would burn. As it burns, it creates a large volume of high energy gas, which is discharged from the motor casing, producing the rocket thrust.

Because the solid fueled rocket motor has no additional systems for pressurized storage, cooling, or pumping of its fuel, it weighs far less than a comparable liquid fueled rocket. The higher mass of the liquid fueled rocket engine reduces its ultimate efficiency below that of the solid fueled rocket, even though the liquid combustion process is chemically more efficient.

For these reasons, it is both desirable and theoretically possible to produce a SSTO using a single solid fuel rocket motor. As a practical matter, this has not been possible using traditional solid rocket motors, however.

Although traditional solid rocket motors can provide the appropriate levels of thrust that would be needed to create a SSTO vehicle, there are difficulties in attempting to use such a motor as the sole engine in an orbital delivery system.

In a traditional solid rocket motor, the solid propellant inside the rocket casing burns along the surface of a hollow inner core within the mass of solid propellant inside the engine. Because of this hollow center core, these designs are known as"center-perforated" (CP) designs. In a CP design, the pressure which is created by the burning of the rocket fuel is directed outward, toward the walls of the rocket casing. As a result, the casing of a CP rocket motor must be quite strong structurally in order to withstand the internal pressures to which it will be subjected when the propellant matrix begins burning. In order to make the casings sufficiently strong, they must be made quite heavy.

This additional weight must then be accelerated all the way to orbit, which in turn requires more thrust.

As the thrust required for the engine goes up, the size of the motor must go up as well. A greater burning surface must be used in order to generate greater thrust, or the burn time must be made longer. This results in CP designs which require still stronger (and therefore heavier) casings to withstand the increased pressures of the larger engine.

Even in circumstances where the payload to be delivered to orbit is sufficiently small that an ordinary sized solid rocket motor can be used, there are still difficulties with attempting to do so. The most significant of these involves maintaining an appropriate acceleration profile. As the motor accelerates the payload toward its orbital altitude and velocity, it burns off fuel, making the vehicle as a whole lighter. Because the vehicle is now lighter, the same amount of thrust being produced by the rocket results in a greater acceleration for the vehicle. Although greater acceleration results in arriving at orbital velocity and altitude more quickly, it can also crush the payload.

In a typical solid rocket motor, the thrust is nearly constant through the entire burn of the engine. As a result, the acceleration provided by this motor will continue to climb as the motor burns off propellant. This problem is actually further compounded by the lightweight nature of the solid rocket motor. Because so much of the motor is propellant, rather than structural mass, the difference between the initial mass of the motor and the final mass of the motor is so extreme that the constant thrust provided can result in an extreme differential. in acceleration. For example, if a Space Shuttle solid rocket booster were launched by itself, its acceleration at burnout would be about ten times as high as its initial acceleration at launch. While some payloads might survive this sort of constantly increasing acceleration profile, most payloads, particularly human crews, will require a lower peak acceleration and a more nearly constant acceleration profile.

One way to achieve a more uniform acceleration profile is to tailor the thrust throughout the burning of the motor so that the thrust gradually reduces as the mass of the vehicle is reduced due to expended propellant. This feature is fairly simple to implement on liquid fueled rockets by varying the rate at which fuel is pumped into the combustion chamber. Traditional solid rockets provide no convenient way to vary their thrust significantly over the course of their burn time. This prevents traditional solid fuel rocket motors from being suitable for use in SSTO vehicles.

An additional related difficulty to the need for variations in thrust profile throughout the launch has to do with the operating point of the rocket engine. Every rocket engine operates ideally at a particular combination of mass flow rate, internal pressure, and exhaust flow expansion. This combination can be referred to as the"operating point" of the motor. Although most rocket motors will operate at conditions other than their ideal operating point, they do so at lower efficiency, and often at extremely lower efficiency.

While steps can be taken to control the factors contributing to the operating point, certain factors necessarily change as the rocket changes from its initial configuration at launch to its final configuration at orbital insertion. One of these is the difference in ambient pressure, which affects both the pressure within the rocket's combustion chamber and the appropriate level of exhaust flow expansion. However, these changes, particularly the

change in chamber pressure, can also lead to changes in burn rate or combustion efficiency, necessitating further adjustments.

As discussed above, liquid fueled rocket engines can be throttled in real time to vary their fuel flow rate, which will allow the vehicle to alter its mass flow and internal pressures. Using these controls, the liquid motor can be maintained within an acceptable range of operation about its ideal operating point. However, traditional solid rocket motors have no such control means available. Once ignited, they remain burning, and their mass flow rates tend to remain substantially constant. As a result, a traditional solid rocket motor will tend to deviate increasingly from its ideal operating point as it burns, resulting in a much less efficient vehicle.

Summary of the Invention In the present invention, an end burning configuration of a solid propellant matrix is used, allowing a varying thrust profile to be applied over the duration of the motor burn. Through the combination of a higher burn rate fuel and an end-burning configuration, flexibility in the structural design of the rocket will enable a single stage, solid fueled motor to accelerate a payload into Earth orbit.

In one preferred embodiment of the present invention, a single-stage-to-orbit vehicle is comprised of a payload, and at least one rocket motor which operates continuously from launch until orbital insertion. The rocket motor comprises a casing, which contains a solid propellant matrix which is burned, creating high pressure exhaust gases which are vented from the motor and produce thrust.

Additionally, the propellant may be comprised of a solid homogeneous mixture of fuel particles which are distributed within a matrix of solid oxidizer. This solid propellant matrix may also be configured so that the burning surface is initially located at the lower, or exhaust end of the rocket, and progresses as it burns toward the top, or payload end, of the rocket.

The propellant matrix may also be constructed so as to provide different levels of thrust at different times during the course of its burn. The propellant matrix may also be shaped so that the size of the exposed burning surface of the propellant will vary over the course of the burn of the motor.

In another preferred embodiment of the present invention, a reusable single-stage-to-orbit payload delivery system is comprised of a transport vehicle, which provides storage for a payload and at least one receptacle for a solid propellant cartridge. In order to launch the system, a solid propellant cartridge is mounted within each receptacle of the transport vehicle. The solid propellant cartridge is ignited and burns, producing exhaust gases, which are vented from the vehicle, producing thrust.

The receptacles may be built into the transport vehicle such that residue from previous cartridges may be removed, and new cartridges mounted within the receptacle without having to substantially disassemble the transport vehicle.

In a further preferred embodiment, the propellant cartridges may be constructed so that they contain only consumable materials and are completely consumed during their use.

Brief Description of the Drawings Figure 1 shows a SSTO vehicle using a solid fueled rocket motor incorporating a variable thrust profile and a modified end-burning configuration.

Figure 2 shows a SSTO vehicle with a solid fueled rocket motor with an end-burning propellant configuration providing a variable thrust profile using a variable casing cross section.

Figure 3 shows a reusable SSTO payload delivery system which uses replaceable solid rocket cartridges to launch the vehicle and lift it into orbit.

Figure 4 shows a"self-contained"solid rocket cartridge which includes a motor casing and a nozzle.

Figure 5 shows a fully consumable solid rocket cartridge.

Detailed Description of the Preferred Embodiment SOLID FUELED MOTOR DESIGN : FIGURE 1 shows a schematic view of a SSTO vehicle making use of a solid fueled rocket motor. The vehicle (20) comprises a payload (40) and a motor (50). Those skilled in the art will recognize that the payload can be any cargo which is to be delivered into or beyond Earth orbit. These can include, but are not limited to, satellites to be deployed, raw materials to be placed in orbit, manned capsules or reentry vehicles, and test articles.

The rocket motor (50) is of a solid-fueled design. A rocket motor casing (60) is used to contain the solid propellant matrix (70) and to attach the motor to the payload (40). Any additional structural systems of the engine are also mounted to the rocket case. These can include, but are not limited to, guidance control means, such as aerodynamic surfaces (80), an exhaust expansion nozzle (82), and ignition means.

The propellant matrix (70) is ignited, and once ignited burns continuously. As the matrix is consumed, combustion by-products are produced in the form of high pressure, high-temperature gases. These exhaust gases (110) are expelled from the rear of the motor, passing through an aperture (100) at the rear of the motor, and into any expansion nozzle which may be used (82). The expulsion of these gases from the motor (50) at high speed produces thrust in the direction opposite which the gases exit the motor. This thrust is used to accelerate the vehicle (20).

PROPELLANT COMPOSITION : The propellant compositions used in accordance with preferred embodiments of the present invention comprise a substantially homogeneous mixture of micron or nanometer-sized particles of metallic fuel particles distributed throughout a matrix of an oxidizer in solid form. A homogeneous mixture, as that term is used herein, means a mixture or blend of components that is generally uniform in structure and composition with little variability throughout the mixture. Different portions of a homogeneous mixture exhibit essentially the same physical and

chemical properties at every point throughout the mixture. The stoichiometry in a homogeneous mixture is also substantially constant throughout the mixture.

Another way of describing the preferred propellant compositions is to say that the metallic fuel and oxidizer are intimately mixed. Intimately mixed, as that term is used herein, means that the two components are present in a structure that is not composed of discrete particles of the two materials, instead the metallic fuel is embedded within a network, crystal, or crystal-like structure of the oxidizer such that the two components cannot be unmixed by general physical methods, e. g. unmixing requires re-solvating or dispersing the oxidizer in a solvent.

In especially preferred embodiments, the propellant comprises a propellant composition called"NRC-3 or NRC-4."Because these two propellant compositions are identical, for purposes of this discussion, they are used interchangeably. In NRC-4, the metallic fuel is aluminum particles having an average diameter of about 40 nm, and the oxidizer is ammonium perchlorate (AP). The aluminum and AP components of NRC-4 are present in stoichiometric quantities, that is, they are present in the quantities needed for reaction, without an excess of any component left over after the reaction.

NRC-4 is preferably made by making a solution of the AP oxidizer in water, and then adding the aluminum particles to the oxidizer solution. The resulting mixture is agitated or otherwise mixed, to produce a substantially homogeneous mixture. The water is then removed from the mixture by freeze drying, as to maintain the homogeneous nature of the mixture, which results in a powdered solid in which the aluminum particles are distributed generally uniformly throughout the solid AP oxidizer matrix. This may also be characterized as controlling the average distance between the metallic fuel particles in the propellant composition.

Example 1 Preparation of AP/Aluminum Nanoparticle Matrix (NRC-3 and NHC-4) Two 50 gram batches of ammonium perchloratel nanoaluminum matrix were sequentially prepared, each by dissolving 25 grams of ammonium perchlorate (0. 5 gram, 99. 9% pure, Alfa Aesar stock #11658) in 0. 5 liters of deionized water to form a solution having a concentration of approximately 0. 4 moleslliter. As in the previous examples, the specific concentration achieved is not critical as long as the solution is well below the saturation point, to ensure that all of the ammonium perchlorate dissolves. To this solution was added 25 grams of nanoaluminum of average particle diameter 40 nm. The quantities of ammonium perchlorate and nanoaluminum were selected so as to yield a stoichiometric ratio of the ammonium perchlorate to the unoxidized aluminum in the nanoaluminum particles.

The mixture was agitated by mechanical shaking to ensure that the particles were completely immersed and that the mixture was substantially homogeneous. The mixture of nanoaluminum particles in ammonium perchlorate solution was then rapidly frozen by pouring the mixture into a container of liquid nitrogen. The container of liquid nitrogen and frozen mixture was then transferred to a vacuum container capable of achieving a base pressure of 10-5 Torr or lower in order to achieve low enough pressure to achieve rapid freeze drying. The vacuum system used was a custom pumping station using a Varian VHS-6 oil diffusion pump, a Leybold-Hereus TRIVAC D30A roughinglbacking pump, and a 16-inch diameter x 18-inch tall stainless-steel bell jar. Active pumping on the vacuum container was immediately

initiated after pouring the agitated mixture into the liquid nitrogen. After a period of 10 minutes, the pressure in the system achieved a steady-state pressure, stabilizing near the equilibrium vapor pressure of the frozen water, i. e., 10-3 Torr. The pressure was maintained at this steady state while the frozen water in the mixture was removed from the mixture by sublimation. After 120 hours removal of the water was complete. It is likely that the time required for water removal can be shortened to some extent by modifying the pouring process to yield a frozen mass of high surface area ; i. e., thin, flat frozen masses as opposed to a single monolithic lump of frozen material. Small, thin frozen masses are expected to dehydrate more quickly during freeze drying than a single, monolithic mass of equivalent weight due to the larger surface area that is exposed by having many small masses relative to the surface area of a single large mass. The resulting processed material of each batch consisted of about 50 grams of low-density, dry agglomerates of particles of ammonium perchloratelnanoaluminum matrix (labeled NRC-3 and NRC-4, respectively).

By changing the size of the aluminum particles used in a propellant composition made according to procedures such as described above, propellants having different performance characteristics may be made. This is because reaction rates, such as the burn rate of a particulate propellant mixture, correspond to the reactant diffusion distance, which corresponds to particle size in particulate materials. Thus, as compared to a propellant using aluminum fuel particles 100 nm in diameter, a propellant using aluminum fuel particles on the order of 30 microns would burn more slowly, release its energy more slowly, and a given mass of propellant would burn over a longer period of time. Conversely, a propellant having aluminum fuel particles of 50 nm would burn faster than the propellant having 100 nm fuel particles, providing greater power in a shorter period of time. Therefore, by choosing the proper size metal fuel particles to include in a propellant composition, a propellant could be made having desired performance characteristics. For the avoidance of doubt, these statements assume that all other things in the propellant, other than particle size, are equivalent.

When changing particle size, one must take the passivation layer into account in order to maintain the correct stoichiometry. When the aluminum is in the form of micron-sized particles, the Al203 passivation layer, which is approximately 2. 5 nm thick, is practically negligible in weight compared to that of the unoxidized metallic aluminum within the particle. However, when the aluminum is in the form of nanometer-sized particles, the aluminum oxide passivation layer can comprise a substantial portion of the total weight of the particle, e. g., 30 to 40 wt. % or more.

Therefore, when nanometer-sized particles are used, less oxidizer per unit weight aluminum fuel is needed for a stoichiometric mixture.

Another way of making a propellant having desired performance characteristics, or of varying the performance characteristics of a particular propellant such as NRC-4, is to make a mixed propellant, comprising at least two fuelloxidizer propellant mixtures. A two component mixed propellant will generally comprise a faster burning propellant component and a slower burning propellant component, at least one of which is a substantially homogeneous mixture of metallic fuel particles distributed throughout a matrix of an oxidizer in solid form, as described above. Additionally, in each of the propellant components, the fuel and oxidizer is preferably present in stoichiometric quantities. The propellant components may have one or more materials in common. For example, a preferred two-

component mixed propellant is one which comprises 200 nm aluminum in a matrix of AP as the faster burning propellant component, and 30 micron aluminum in a matrix of AP as the slower burning propellant component. Another preferred two-component mixed propellant is that which comprises 85% by weight of NRC-4 as the faster burning propellant component and 15% by weight of the slower burning propellant component comprising hydroxy-terminated polybutadiene (HTPB) and AP in stoichiometric quantities. However, any fuelloxidizer propellant may be used, and mixed propellants may contain more than two propellant components.

When a propellant formulation comprises two propellant components, a faster burning propellant component and a slower burning propellant component, it will burn at a rate that is dramatically limited by the burn rate of the slower burning propellant component. If the burn rate of both components is known, the amount of each component needed to create a propellant of a desired burn rate may be approximated by using Equation 2 : <BR> <BR> <BR> <BR> (. +)<BR> <BR> <BR> R = mmrU t- (Eq. 2)<BR> <BR> <BR> mf IRf +m, IR, wherein R is the desired burn rate, mS is the mass of the slower burning propellant component, m, is the mass of the faster burning propellant component, Rs is the burn rate of the slower burning propellant component, and R, is the burn rate of the faster burning propellant component. Although these burn rates are in terms of mass per unit time, burn rate may also be expressed in terms of length per unit time as in the data presented herein. Because Equation 2 is based upon several assumptions, the results regarding observe rates or needed quantities may vary slightly from the calculated values. In some circumstances, it may be desirable to optimize the formulation calculated using the equation above. Optimization may be done experimentally by preparing mixed propellants and testing them in the laboratory or in the field. By using a relation such as Equation 2 andlor the principles embodied therein, the burn rate characteristics of a mixed propellant can be"tuned"to fit a particular application or use, dependent upon the amount of propellant components added and the difference in burn rate between the faster and slower burning propellant components.

For presently preferred applications, HTPBIAP is used as the slower burning propellant component due to its low cost, availability, and well-understood properties. However in some motors it may be desirable to use a slower burning propellant component having a burn rate faster than that of HTPBIAP, i. e. one having a burn rate closer to the faster burning propellant component. One advantage in using such materials is that it is easier to fine tune the mixed propellant and to manufacture consistent batches of mixed propellant, because each gram of HTPBIAP propellant has a higher net effect than each gram of a slower burning propellant component having a burn rate faster than HTPBIAP, as can be demonstrated using Equation 2. For example, one may substitute a homogeneous mixture of 30 micron aluminum particles in a matrix of AP for HTPBIAP as the slower burning propellant component when used with NRC-4 as the faster burning propellant component. Because the propellant comprising 30 micron aluminum as the fuel is closer to the burn rate of NRC-4 than a propellant having HTPB as the fuel, relatively small changes in composition will result in smaller changes in overall mixed propellant performance.

The two or more components in a mixed propellant are preferably mixed together to achieve a substantially consistent, well-mixed mixture. Such a mixture of components in the mixed propellant helps to avoid having uneven burn rates, power or other properties in large portions of the propellant bulk. If one or more components are present in a quantity or form that makes it difficult to achieve consistent mixing or a consistent composition in the mixture, one may achieve a well-mixed propellant by use of a solvent. In using a solvent to aid mixing, one combines the various components of the propellant in the solvent, mixes the resulting mixture by agitation, stirring, sonicating, etc. to form a solutionisuspension, and then removes the solvent. A solvent used to aid mixing is chosen for its compatibility with one or more of the components of the mixture, such as miscibility with a component or ability to dissolve a component.

Preferred solvents will not substantially react with the fuel, oxidizer, or other components of the propellant mixture. For propellant compositions comprising aluminum, AP and HTPB, such as the preferred mixed propellant composition disclosed above, preferred solvents include nonpolar solvents such as hexane or pentane. Because the solvent is removed by evaporation, such as in open air, under reduced pressure, with application of heat or other method as is known in the art, solvents having a low boiling point or high vapor pressure are preferred.

Example 2 Preparation of Propellant Mixture A small-scale, 1-gram batch of propellant was prepared by dissolving 0. 047 gram of HTPB into 15 ml of reagent grade hexane in a capped, cylindrical glass container of approximately 25 ml volume. To this solution, 0. 103 gram of AP (3-micrometer particle size) was added, followed by 0. 85 gram of NRC-3. The resulting mixture was sonically mixed for about 10 minutes. The hexane was removed by evaporation in air with warming to about 40 C, to leave a solid propellant material.

It is well known in the propellant industry that propellants generally burn faster at higher pressure. The behavior is usually described by the formula where Rb is the burn rate, C is a constant, P is pressure, and n is the pressure exponent. It is further widely known in the industry that the value of the pressure exponent for a candidate propellant is critical to the utility of the propellant in rocket motors. In particular, if the value of the pressure exponent for a candidate propellant is 1 or greater, the candidate propellant is unsuitable as a rocket propellant, as the burn rate will increase uncontrollably as pressure builds and will thus lead to an explosion. On the other hand, if the exponent is 0. 6 or lower, the candidate propellant will be relatively stable in typical rocket motor environments.

The burn rate and pressure exponent of the propellant produced in Example 2 was determined by measuring the burn rate at high density at various pressures by pressing the propellant into pellets and measuring the burn rate in a sealed pressure vessel at various applied pressures. Several high-density pellets were formed from the propellant mixture of Example 2 by pressing nominally 0. 080 grams of the propellant mixture for each pellet into a cylindrical volume measuring 0. 189 inches in diameter and approximately 0. 1 inches long, using a hydraulic press and stainless

steel die assembly. A density of approximately 1. 7 grams per cubic centimeter was obtained by applying a force of 400 pounds to the die. A free-standing, cylindrical pellet, thus formed, was removed from the die by pushing the pellet out of the die.

The burn rate of a free-standing pellet can be measured by burning the pellet in a confined volume and measuring the pressure rise as a function of time in the volume. As the pellet burns, the product gases formed by the propellant will cause the pressure in the confined volume to increase until the burn is complet. By measuring the length of the pellet before the burn and measuring the time interval during which the pressure increases during the burn in such a volume, the average burn rate of the propellant can be calculated by dividing the pellet length by the time interval that the pressure was increasing. Performing such measurements with the confined volume pre-pressurized with a non-reactive gas (e. g., dry nitrogen) yields burn rates at eievated pressures that can be used to calculate the pressure exponent for the propellant.

Example 3 Burn Rate Testing and Pressure Exponent Determination of Propellant Mixture Three pellets fabricated from the powder prepared in Example 2, as described above, were separately burned in a stainless steel pressure vessel of 350 cubic centimeters, to determine burn rate and the burn rate exponent for the propellant mixture. The pressure vessel contained a pressure transducer (Endevco, 500 psig) and two electrical connectors to which a hot wire igniter (nichrome wire, 3 inches long by 0. 005 inches in diameter) was attached. In each of separate tests, the igniter wire was first taped to the flat bottom of the pellet, the igniter wire (with pellet) was attached to the electrical connectors inside the pressure vessel, and the vessel was sealed. The pellet was ignited by passing a 3-amp DC current through the electrical connectors, causing the igniter wire to heat and ignite the propellant. Pressure in the vessel was recorded as a function of time by measuring the electrical output of the pressure transducer with a digital oscilloscope (Tektronix, model TDS460A). One of the pellets was burned at the ambient atmospheric pressure of the laboratory. The other two pellets were burned after pre-pressurizing the vessels with dry nitrogen to 125 and 300 pounds per square inch, respectively. Pellet weight, pellet length, pellet density, burn time, and average pressure during the burn for the three pellets are shown in Table 1.

Table 1 lligh-DensitJ/BUM Rate Results Weight (g) Length (in.) Density (gicc) 1 Time (sec) Burn Rate Pressure (psig) (inlsec) I 0. 060 0. 080 1. 63 0. 0286 2. 80 16. 6 0. 080 0. 107 1. 63 0. 0132 8. 11 167. 5 0. 085. 112 1. 65 0. 0111 10. 08 338. 1 A least-squares polynomial fit of the data in Table 1 reveals that the burn rate for this propellant varies as Rb = (0. 8374) p (0 4337>,

Where Rb is burn rate in inches per second and P is pressure in pounds per square inch. The pressure exponent, n, for this propellant mixture is approximately 0. 43 (i. e., n < 0. 6), suggesting the mixture should be acceptable for rocket motor applications, from a pressure-dependence perspective.

In order to compare propellant formulations of the present invention, both to each other and to the prior art, a simple laboratory scale test was devised. The propellant compositions tested were made according to the solvent- based method described above. The test allows for the measurement of properties relevant to the performance of a propellant, such as burn rate, average thrust, and Propulsion Potential (lisp at very low, near ambient pressures). The test provides for the measurement of weight (force) and time while the propellant is being burned in a mini-motor.

Because some properties may be dependent in part upon factors including the size andlor aspect ratio of the motor, particular motor configurations were chosen for use in the tests. One configuration chosen for the mini-motor was a stainless steel tube having an internal diameter of 0. 19 inches and an aspect ratio of about 12 : 1 (length to internal diameter). Another series of tests were done using the same 0. 19 inch ID stainless steel tubing in which the aspect ratio was about 5 : 1.

To perform the test, a section of the 0. 19 inch ID stainless steel tubing was cut to a length (within about 5%) to provide a motor having the desired aspect ratio for that series of tests, and filled with propellant to make the motor. The filling was done by placing the propellant into the tube, and then tamping or packing it down into the tube, first by hand and then by means of a laboratory press. A sleeve was placed on the tube to provide balance and support, which was then placed on an electronic balance and zeroed. The motor was then ignited and the mass or force, in grams, was measured as a function of time. From these data points, the mass of propellant, burn time, burn rate average thrust and Propulsion Potential were be calculated.

The tests comparing two NRC-4 formulations to three more conventional propellant formulations were performed as discussed above, and used mini-motors having an aspect ratio of approximately 5 : 1 (length to internal diameter). The results of the tests are set forth in Tables 2 and 3 below.

Table 2 NHC-4 Propellants in the 5. 1 Mini-Motor Composition Propellant Burn rate Burn Time Average Propulsion (g) (inisec) (sec) Thrust (g) Potential (sec) (Isp) 1 65% NRC-4 ; 11. 1% 0. 574 0. 395 1. 98 5. 814 20. 1 HTPB ; 23. 9% 3p AP 2 60% NRC-4 ; 12. 6% 0. 564 0. 373 1. 86 5. 901 19. 5 HTPB ; 27. 4% 3 AP I I I Table 3 Conventional Propellants in the 5.-1 Mini-Motoi (nn intimate mixing of Al/APl Composition Propellant Burn rate Burn Time Average Thrust Propulsion (g) (inisec) (sec) (g) Potential (sec) (lsp) 3 19% 30L Al ; 69% 2001l 0. 935 0. 030 38. 56 0. 025 1. 0 AP ; 12% HTPB 4 19%5 ; 69%3AP ; 0. 662 0. 059 17. 52 0. 057 1. 5 12% HTPB 5 19% 3 au ; 69% 3p AP, 0. 630 0. 064 15. 82 0. 098 2. 5 12% HTPB

Much of the discussion presented herein is in terms of burn rate. This is because the burn rate of a material is highly indicative of its properties and suitability as a propellant. However, for experimental purposes, one generally uses the specific impulse (Isp) for comparison. The Isp takes the amount of the propellant material tested into account, thus allowing for a direct comparison between the various formulations and tests for which there may be slight differences in the quantity of the material used.

It should be noted herein that the data presented in Tables 2 through 5 for the propellant formulations are values that were measured when the propellant was combusted under a very low, near ambient pressure. No nozzle or other flow restrictor was placed on the tubes during burning, nor was there any other method used to increase the pressure of the material during combustion. This differs from the general practice in the aerospace industry, wherein lsp values are generally measured at a pressure of 1000 psi and reported as such, oftentimes without indication that such elevated pressure was used. If the pressure is increased, one expects the burn rate to increase, which would lead to an increase in measured Isp due to the relation between the two properties. Therefore, in the discussion which follows the measured Isp at near-ambient pressures will be termed"Propulsion Potential"to avoid confusion with and distinguish from the industry-standard high pressure Isp measurements.

Table 2 presents the results of tests on two propellant formulations of the present invention using NRC-4 powder. The amount of AP listed in the composition is the stoichiometric amount of AP for the HTPB present, that is the amount of AP needed to react the HTPB only. The NRC-4, as discussed supra includes AP in a quantity sufficient to react with all the aluminum component thereof. Table 3 presents the results of tests on three more conventional propellant formulations in which the components as listed are micron-sized and are mixed together and cast into the tubes without curing. The AP listed in the formulations of Table 3 is the stoichiometric amount for both the AI and HTPB present. The formulations in Table 3 do not comprise the intimate, homogeneous mixtures of aluminum and AP of the compositions of the present invention, including NRC 4. All compositions in both tables, however, have about 12% HTPB. All percentages herein are by weight.

The results of Table 3 demonstrate the effect of particle size, and thus reactant diffusion distance, as discussed herein. Formulation 3, comprising 30µ AI and 200 AP has the largest particle sizes, followed by formulation 4 having 5µ AI and 3µ AP, and finally by formulation 5 having 3µ Al and 3µ AP. It can be seen from

Table 3 that the Propulsion Potential increases as the particle size decreases, indicating that the lower particle size formulations would provide more powerful fuels.

An additional factor which may be at work is the difference in the particle sizes. In formulation 3, the AP particles are, on the average, about 6-7 times larger than the Al particles. In formulation 5, the particles of AI and AP have the same average diameter. The size difference between the particles in formulation 3 would make homogeneous mixing of the fuel and its oxidizer difficult, which could also, or alternatively, account for its lower Propulsion Potential and lower burn rate.

Comparison of the data in Table 2 to formulation 5 in Table 3 shows that the Propulsion Potential is increased about 8-fold when the fuel and its oxidizer is in the form of an intimate, substantially homogeneous mixture of nanoaluminum and AP according to a preferred embodiment (NRC-4) of the present invention. In these formulations, the NRC-4 provides small fuel particle size, on the order of about 40 nm, as well as low reaction diffusion distance because the nanoaluminum is dispersed throughout the AP oxidizer phase in a substantially uniform fashion. In preferred embodiments of fuelloxidizer matrix compositions, such as NRC-4 and similar compositions comprising larger, micron-size fuel particles, the concerns regarding obtaining a homogeneous mixture of fuel and oxidizer seen in formulation 3 are minimized, because the composition itself, having the fuel particles dispersed throughout the oxidizer phase provide a mixture which is substantially homogeneous, intimate, and of the correct stoichiometry.

Thus, it can be seen that the preferred propellants have very high energy, power, and burn rate as compared to propellants comprising more standard-like particle mixes.

Several additional mixed propellants, comprising two components (i. e. propellants, fuelloxidizer mixture), have been prepared, and tested according to the general procedure described above. The propellants made had varying amounts of low and high burning rate propellant components. The composition is listed in the tables in terms of the quantity of NRC-4 present, expressed as a percentage by weight. The remainder of the propellant comprises HTPB and its stoichiometric quantity of AP. The mixed propellants were made by mixing the various components together in the presence of nonpolar solvent which is later evaporated, as described above. The HTPB in the propellant formulations was used neat, without a curing agent, such that the propellant could be loaded into the test motor immediately after mixing and burned thereafter, without having to wait for the material to cure, although it was not a necessity that the loading and testing be done immediately following mixing. Additionally, burn rate catalyst was not added to the propellant mixtures tested herein. The results of these experiments are presented in Tables 4 and 5 below.

Table 4 NRC-4 Containing Propellants in the 12. 1 Mini-Moto % NRC-4 Propellant Burn rate Burn Time Average Thrust Propulsion (g) (inisec) (sec) (g) Potential (sec) (lsp) 70 1. 519 0. 933 1. 59 30. 527 31. 9 60 1. 411 0. 434 4. 56 35. 626 25. 2 50 1. 770 0. 250 8. 57 1. 888 9. 1

Table 5 NRC-4 Containing Propellants in the 5:1 Mini-Motor % NRC-4 t Propellant Burn rate BurnTime | AverageThrust Propulsion (g) (inisec) (sec) (g) Potential (sec) (lsp) 65 0. 574 0. 395 1. 98 5. 814 1 20. 1 60 0. 564 0. 373 1. 86 5. 901 19. 5 50 0. 443 0. 361 1. 97 2. 041 9. 1 40 0. 537 0. 182 5. 22 0. 403 3. 9 35 0. 568 0. 139 7. 19 0. 265 3. 4 20 0. 615 0. 056 19. 17 0. 053 1. 7 As can been seen in the tables above, relatively small changes in the composition of the propellant (ratio of high and low burn-rate components) can have a dramatic effect on the Propulsion Potential when the propellant is combusted. Furthermore, tests such as those above can be used to aid in devising a formulation to achieve particular results. Using the data above, for example, if one wanted to make a propellant having a Propulsion Potential of 5, one would need to prepare a propellant having a little over 40% NRC-4 by weight if a 5 : 1 mini motor were used. The formulation required may be found more exactly by methods known in the art, including fitting the experimental data to an equation or iteratively by preparing and testing additional formulations within the narrowed ranges determined using the data above.

The results of additional experiments conducted by the Inventors are presented in Appendix 1 hereto. These tests were conducted using laboratory scale mini-motors of varying aspect ratios, some of which also comprised a flow-restricting nozzle. Appendix 1 details the formulation (% NRC-314 to % HTPB with its stoichiometric quantity of AP), the mass of the propellant in grams, the density at which the propellant is packed in the motor casing, the pressure in the combustion chamber, whether there was a nozzle present, the orifice size of the nozzle, the length of propellant in the motor casing, the burn time, the burn rate, the aspect ratio, the thrust, and the Isp for several different mixed propellant compositions. The blank spaces indicate where particular data is unavailable or not applicable.

This data and the other information set forth herein support the proposition that reasonable thrust is achievable at lower pressures. While a typical thrust analysis of a conventional rocket motor involves a high pressure

component, one should realize that this higher pressure at which combustion occurs is not achieved without a loss of energy in the exhaust gases. That is, such higher pressures are typically achieved by means of throat or a nozzle which"chokes"the flow of the exhaust gases. True, such a nozzle increases the speed of the gases through the nozzle but it also decreases the energy of other gases which impinge on the narrowed throat structure. This in turn results in an increased pressure which heretofore has been necessary to increase the burn rate.

However, given a chemical reaction which produces sufficient energy and higher burn rates at lower, say near ambient pressures, there is no reason why reasonable thrust cannot be achieved without a nozzle and the associated higher pressure. In other words, the kinetic energy of the combustion, which produces expanding gases having a given mass moving at a high velocity, is sufficient to produce the momentum transfer necessary to achieve reasonable thrust. This is achieved in the present case by relatively high burn rates at near ambient pressures, which burn rates were not previously achievable without higher pressures. Of course, at higher pressures which could be achieved with some type of throat or nozzle device, even higher burn rates are likely to be achievable. Thus, rocket motors utilizing propellants of the type described herein operating at pressures other than ambient or near ambient are also within the scope of the preferred embodiments.

Additional details not necessary to repeat here are disclosed in assignee's copending applications entitled COMPOSITION AND METHOD FOR PREPARING OXIDIZER MATRIX CONTAINING DISPERSED METAL PARTICLES, application Serial No. 091447, 703, and VARIABLE BURN-RATE PROPELLANT, application Serial No. 091448, 546, filed on the same date as the present application, the entireties of which is hereby incorporated by reference.

END-BURNING CONFIGURATION : Once the propellant matrix (70) is ignited, it will burn at its exposed surface (90). This is referred to as the "burn-front". As fuel burns away, the burn-front will progress farther into the propellant matrix, moving in a direction substantially perpendicular to the surface of the burn-front itself. As a result, the burn-front will spread into the casing (60) of the rocket as the propellant continues to be consumed in the combustion process.

In accordance with a preferred embodiment of the present invention, the propellant matrix is constructed into a configuration in which the burn-front (90) will progress in a direction substantially toward the payload (40) at the front of the vehicle. Such a situation where the burn-front (90) is located at the rearmost portion of the propellant matrix (70) is referred to as an"end-burning"configuration. This is a departure from the traditional solid rocket motor configuration, in which the propellant matrix is designed with a centrally located hollow space at which burning is initiated. In these"center-perforated" (CP) designs, the burn front progresses substantially from the center of the propellant matrix toward the sides of the rocket casing (60).

Using traditional solid rocket fuels, creating an end-burning design is problematic because the burn-front is smaller in surface area than in a similarly sized CP design. Because of the slower burn rate found in traditional fuels, the greater surface area is necessary in order to have a sufficient mass flow rate in the rocket to produce the desired

thrust. Furthermore, the additional mass flow is needed in order to maintain a sufficient amount of pressure on the burn-front to maintain effective combustion.

Using the high burn rate solid propellant discussed above, it becomes feasible to construct a solid rocket motor in an end-burning configuration and still generate sufficient thrust to perform useful launch operations. In addition to the high burn rate properties of the NRC-4 and similar fuels, its ability to sustain combustion at ambient pressure allows for a wider range of operational conditions with respect to the pressure found in the chamber.

The end-burning configuration also reduces the loads imposed upon the rocket casing (60), and therefore reduces the strength needed for its design. This is because the pressure produced by the burning propellant is lower than in a comparable CP design for a solid rocket motor, and so the case need not support the same degree of internal loading that would be required in a CP design.

Because NRC-4 and similar propellants provide sufficient operating thrust even when operating at or near ambient levels of pressure, there is no need to increase the level of pressure within the rocket casing beyond what occurs due to the burning of the fuel itself. By dispensing with the need for additional back-pressure, the pressure within the rocket casing is lowered, and casing need not be designed to contain this additional pressure.

In addition, the end-burning design allows the propellant matrix to be constructed in such a way that it can provide its own structural support to a greater degree than in comparable CP propellant matrix designs. Because the propellant matrix is more stable structural, the case need not provide as much structural support to the propellant as would be required in a CP design.

These reduced strength requirements for the propellant casing broaden the range of materials and configurations available for the propellant casing, as well as the manner in which the propellant can be mounted within the vehicle itself. Instead of steel or high strength composites, the casing may be constructed from lighter, less strong materials. These may include, but are not limited to, plastics or ceramics. Additional details not necessary to repeat here are disclosed in assignee's copending application entitled END-BURNING ROCKET MOTOR, application Serial No.

091447, 758, filed on the same date as the present application, the entirety of which is hereby incorporated by reference.

The use of lighter weight materials in the construction helps reduce the overall weight of the vehicle, which increases its effective fuel fraction. By increasing the fuel fraction, it becomes more feasible for a single stage vehicle to possess sufficient total thrust to boost a payload into orbit.

VARIABLE THRUST PROFILE PROPELLANT MATRIX : An additional important benefit of using an end-burning configuration is that it provides the capability to produce a variable thrust profile motor, while retaining the solid fuel design. There are several ways that this can be done using an end-burning motor with a high burn rate fuel.

In one preferred embodiment the thrust profile is varied by varying the burn rate of the propellant matrix itself. By changing the density, chemical composition, or particulate size as discussed above, it is possible to alter the

burn rate, and therefore to alter the total amount of thrust that the motor is providing at any given time. Using a traditional solid rocket fuel, the burn rate is not substantially variable. The propellant will tend to burn at a constant rate. Therefore, as long as the burn surface does not change size significantly, the total thrust will not change significantly.

In another preferred embodiment the thrust profile is varied by varying the size of the burn-front at different times during the burn. This is accomplished by choosing a geometry for the rocket casing and propellant matrix such that as the burn-front progresses through the propellant, it becomes larger or smaller based on the shape of the casing and the direction in which the burn-front is oriented. This can be seen in FIGURE 2. Because the burn-front will move substantially perpendicular to its surface at any given point, it becomes possible to design a geometry where the burn- front will move toward the rocket casing (reducing the overall burn-front size) or away from the rocket casing (increasing the overall burn-front size). A similar effect is produced by varying the cross sectional area of the motor itself at different points along its length. If the burn-front is designed to progress in a purely end-burning manner from the exhaust (100) end toward the payload (40) end, a continually tapering profile for the casing (60') will result in progressively smaller cross-sections, and hence burn-fronts (90, 90', 90"), over the duration of the operation of the motor.

Although similar techniques can theoretically be applied in center-perforated rocket motor designs, any attempt to do so encounters practical limitations which render them unsuitable for the design of a solid fuel powered SSTO. To a large degree this is a matter of geometric limitations.

The burn-front of a CP design can be made to decrease from its ignition to its burnout by using a star-shaped design. The more points that the star is given, and the deeper those points penetrate into the propellant matrix, the greater the initial surface area will be. As the propellant matrix is consumed, the star will smooth out into an increasingly circular cross-section, which has a lower surface area than that of the initial star-shaped core.

Theoretically, this allows the surface area to decrease by any desired amount between ignition and burnout.

However, the greater the differential between initial and final surface areas is made, the more certain structural difficulties come into play. By increasing the number of points on the"star", the width of propellant matrix between the points of the star is decreased. This decreased width means a reduced overall burning time before the star collapses upon itself. Increasingly complex interior geometries also make for a more structural weak matrix, more susceptible to cracking and deformation, either of which can result in unanticipated deviations from the desired burn profile, as well as over-pressures due to increased burning surfaces, and the expulsion of unburned propellant from the motor.

Furthermore, increasing the surface area of the internal core in a CP design in this way increases the size of the internal hollow space. This results in an inefficiency in the motor, because now a larger casing is required to contain the same amount of propellant. Complex internal geometries are also difficult to fabricate in CP configurations, raising the cost associated with producing such a design.

The end-burning configuration avoids these difficulties by allowing the size of the burning surface to be controlled by the width of the casing itself. By structuring the surface of the propellant matrix such that the burn front remains perpendicular to the forward direction of the vehicle, but varying the diameter of the rocket casing along its length, it becomes possible to simply design an end-burning motor in which the cross section tapers from the rear to the front. As the burn-front progresses upward along the length of the motor, it will continually shrink. By designing a tapering profile that accounts for the steadily decreasing weight of the vehicle as fuel burns off, it is possible to maintain a constant acceleration of the vehicle while still using a solid rocket motor.

Additionally, one may configure the propellant matrix so that the burn-front evolves from a partial conical, or other angled shape, into a more purely end-burning configuration as the burn progresses. Since the surface of the cone and the flat configurations will be different, not only the size, but the shape of the burn-front will evolve during the operation of the motor. Such a configuration is referred to as a"modified end-burning"configuration, and an example of such a configuration is provided in FIGURE 1.

Since the end burning design does not require a hollow core, none of this central space within the motor is wasted, making for a more efficient use of the space. The lack of deeply cut shapes in the propellant matrix and other structural complex internal geometry also eliminates concerns regarding the structural integrity of the propellant itself.

Through the combination of a high burn-rate fuel and an end-burning design, the needed total thrust and the desired variations in thrust profile can be achieved using a single solid fueled rocket motor, thereby creating a SSTO using a solid fuel. Those skilled in the art will appreciate that the usage of such techniques to create a SSTO solid fueled vehicle is not limited to the application of only one of the above techniques at a time. Using multiple techniques described herein in combination with each other or with other traditional techniques of rocket design is understood to be a way of practicing the described invention.

REUSABLE SSTO PAYLOAD DELIVERY SYSTEM DESIGN : In a further preferred embodiment, self-contained solid fueled motor"cartridges"are used to launch a reusable launch and reentry vehicle. This vehicle can then be refueled by simply inserting new solid propellant cartridges. This is not possible using current solid fuel rocket technology in which the propellant must be cured once in its final position, i. e. inside the casing of the rocket vehicle. By using an end burning design, it is possible to prepare and cure the solid propellant matrix independent of the vehicle in which it will be used, creating a fuel cartridge, which can then simply be loaded into appropriate mounting points in or on the reusable launch and reentry vehicle. A schematic representation of the design of such a system can be found in FIGURE 3.

The payload delivery system (200) comprises a transport vehicle (210) and propellant cartridges (250). The transport vehicle (210) is designed to accept payloads for deployment, which are mounted in a payload storage area (220). Those skilled in the art will understand that the payload storage area need not be internal to the transport vehicle as shown in FIGURE 3, but could include external means for attaching payloads to the transport vehicle, or

other containment and deployment means. The transport vehicle (210) also includes one or more receptacles (240) which are designed to accept propellant cartridges (250). Another feature of the transport vehicle (210) is one or more exhaust apertures (230).

The receptacles (240) are designed so that it is possible for a propellant cartridge (250) to be placed into the receptacle without significantly dismantling the transport vehicle. It will be understood by those skilled in the art that the receptacles need not be fully contained within the transport vehicle. The invention may also be practiced using receptacles which are completely or partially external to the transport vehicle. Furthermore, such mounting and dismounting may occur in a variety of ways, such as inserting the propellant cartridge (250) through the exhaust opening (230), or inserting the cartridge through an alternate opening into the receptacle designed specifically for this purpose.

The propellant cartridges (250) are comprised of a solid propellant matrix, as discussed earlier, and means for mounting the cartridge within the receptacle, and optionally certain other components. Propellant cartridges will be discussed in more detail below in reference to FIGURE 4 (SELF CONTAINED PROPELLANT CARTRIDGE) and FIGURE 5 (FULLY CONSUMABLE PROPELLANT CARTRIDGE).

By allowing the propellant cartridges (250) to be mounted or dismounted from the appropriate receptacles (240) without significantly dismantling the transport vehicle (210), it becomes possible to rapidly refit the transport vehicle so as to use it again relatively soon after a previous use.

REUSABLE SSTO PAYLOAD DELIVERY SYSTEM OPERATION : Prior to launch, propellant cartridges (250) are mounted within each receptacle (240) of the transport vehicle (210). The payload is mounted in the vehicle, and then the vehicle is prepared for launch.

At launch, each propellant cartridge is ignited, and proceeds to burn in the manner discussed above. In operation, the propellant cartridge acts substantially similarly to the solid rocket motors described in previous preferred embodiments. The propellant matrix will burn, producing exhaust gases, which will be vented, producing thrust. In the instant embodiment, the exhaust gases will be directed toward the exhaust aperture or apertures (230) of the transport vehicle (210), where they will exhaust from the vehicle and thrust the vehicle in the opposite direction.

The cartridges may be constructed using any of the techniques described earlier, including end-burning configurations, variable thrust profiles, variable cross sections, and high-burn rate propellants, such as NRC-4. After ignition, the propellant will burn until exhausted, at which point the vehicle shall have achieved its intended altitude and velocity.

After completing its mission and returning to earth, the transportation vehicle (210) will be refitted. This will involve removing any remaining portion of the propellant cartridges (250) from the receptacles (240), mounting new cartridges in the receptacles, and mounting a new payload. At this point, the vehicle is ready to be prepared for another launch.

As can be seen, the transport vehicle (210) is used for multiple launches ; only the propellant cartridges are replace after each use. Furthermore, because mounting the cartridges into the vehicle requires neither dismantling the vehicle nor constructing the cartridges inside the transport vehicle, the turn around time can be significantly shorter than is currently possible using solid fueled rockets.

In existing solid fueled rocket designs, the solid fueled rocket motor is generally disposed of after each use.

In the case of solid rocket motors designed for reuse, such as the Space Shuttle solid rocket boosters, the boosters must be disassembled, refitted, and then completely rebuilt and refilled with propellant after each launch. By allowing the propellant cartridges to be built and stored separately from the vehicle in which they operate, it becomes possible to have a stockpile of cartridges available, allowing for shorter intervals between successive lunches.

SELF CONTAINED PROPELLANT CARTRIDGE : An example of a solid rocket propellant cartridge for use in a preferred embodiment as described above is shown in FIGURE 4. This propellant cartridge (300) comprises a complete solid rocket motor which is mounted to the transportation vehicle (210) of FIGURE 3. Because it is a full motor, the cartridge is referred to as a"self contained" solid rocket propellant cartridge.

The self contained cartridge (300) comprises a casing (310), a solid propellant matrix (320), and attachment means (330) used to mount the cartridge to the transport vehicle (210). The propellant matrix (320) is contained within the casing (310), and is structured as is described above in order to provide the appropriate thrust necessary to boost the transport vehicle into Earth orbit.

When ignited, the propellant matrix (320) will burn at its exposed surface, or burn-front (340), producing exhaust gases which are vented from the rear of the cartridge. The exhaust gases may be vented directly into the ambient environment, or they may be passed through a nozzle (350) which expands the gases before venting them to the ambient environment. It is also possible that the exhaust from the cartridge passes into a chamber within the transport vehicle before being expelled from the vehicle.

For example, one embodiment uses exposed propellant cartridges which vent directly to the ambient environment with a nozzle. Another uses fully contained cartridges which vent into a nozzle which is part of the transport vehicle. Still another embodiment uses an internal cartridge with its own nozzle where the nozzle is exposed directly to the ambient environment. In another, multiple cartridges are exhausted internally to a common nozzle mounted on the transport vehicle. Those skilled in the art will understand that the described invention includes, but is not limited to, each of the above designs.

As the propellant cartridge burns after ignition, the propellant matrix (320) will be consumed. When the propellant is fully consumed, the cartridge will produce no more thrust and is completely expended. When the transport vehicle is refitted for a new launch, the spent cartridge must first be removed. This spent cartridge will still include a casing (310) and attachment means (330), and may include a nozzle (350). Once this spent cartridge is

removed, it may be refitted and refilled for a future launch.. A fresh cartridge containing a new propellant matrix is then installe into the transport vehicle, and the system is prepared for another launch.

By using a durable casing, the attachment means may be mounted to the casing making possible the use of standardized mounting systems, even between vehicles which make use of different size or different thrust cartridges.

This system also allows the nozzle to be matched to the propellant matrix which is being used. This differs from the cartridge described below.

FULLY CONSUMABLE PROPELLANT CARTRIDGE : A solid fuel rocket cartridge used in a different preferred embodiment of this invention is shown in FIGURE 5. This is the fully consumable propellant cartridge (400). This cartridge comprises, at a minimum, a propellant matrix (420) and attachment means (430) for mounting the cartridge within the appropriate receptacle (240) on the transport vehicle (210).

The consumable propellant cartridge may also comprise a casing (410) surrounding the propellant matrix (420). The casing should be constructed from a material such that it will be consumed when exposed to the temperatures and pressures which occur during the burning of the propellant. The casing may even be constructed from a second propellant matrix which has a slower burn rate than the first propellant matrix (420).

This cartridge operates in a manner substantially similar to the self contained propellant cartridge described above. However, the difference is that at the end of the cycle, there is no need to remove the existing cartridge from the receptacle of the transport vehicle. This is because the entire cartridge is constructed using materials which are consumed during the launching of the transport vehicle.

Using a consumable cartridge provides advantages in weight, as well as simplicity of operation during refitting of the transport vehicle. Consumable cartridges of necessity will not include a nozzie on the cartridge, but may make use of a nozzle which is integral to the transport vehicle.

Additional details not necessary to repeat here are disclosed in assignee's copending application entitled NOZZLELESS ROCKET MOTOR, application Serial No. 091447, 757, filed on the same date as the present application, the entirety of which is hereby incorporated by reference.

It will be understood by those skilled in the art that propellant cartridges may be designed which incorporate some aspects of the self-contained cartridge, and some of the fully consumable cartridge. However, those skilled in the art will recognize that the described invention may be practiced using different combinations of the techniques described above. For example, a fully consumable propellant cartridge may be used in a transport vehicle which does not provide its own nozzle. Using multiple techniques described herein, in combination with each other or with other traditional techniques of rocket design, in order to create a SSTO payload delivery system is understood to be a way of practicing the described invention.

APPENDIX 1<BR> Additional Mini-Motor Data %NRC3/47 Prop. Mass Density Pressure Nozzle Nozzle Motor Length Burn rime Burn Rate aspect Trust isp Experlment Run / File %HTPB+AP grams glcc psig Y/N Orifice (in.) prop., in. sec in./sec. Ratio grams sec. scope89.mac/2 60/40 0.84 1.903241 75.1 Y 0.081 0.96 0.683 1.41 5.079365 83.4 67.8 scope87.mac/5 85/15 0.8 1.72801 29.45 Y(.052) 0.081 1.007 0.124 8.12 5.328042 528 81.8 scope83.mac/10 85/15 0.38 1.707749 235.3 Y 0.081 0.484 0.078 6.19 2.560847 477.5 98.0 scope79.mac/13 85/15 0.36 1.544473 173.6 Y 0.089 0.507 0.0947 5.35 2.68254 395.7 104.1 scope77.mac/15 85/15 0.36 1.594802 7.5 Y 0.101 0.491 0.139 3.53 2.597884 214 82.6 scope75.mac/17 85/15 0.37 1.599998 15.6 Y 0.128 0.503 0.18 2.76 2.661376 137.5 66.9 scope73.mac/19 85/15 0.35 1.572926 Y 0.154 0.484 0.294 1.65 2.560847 43 36.1 scope71.mac/21 85/15 0.35 1.586034 Y 0.169 0.48 0.273 1.76 2.539683 44.3 34.6 scope59a-h. dat/31 85/15 0.523 1.702986 N 0.668 0.85 0.79 3.534392 16 26.0 scope58a-h. dat/32 85/15 0.591 1.86647 N 0.685 1.35 0.5 3.624339 6 13.7 scope59a-h.dat/35 85/15 0.523 1.702986 N 0.668 0.85 0.79 3.534392 16 26.0 scope58a-f.dat/37 85/15 0.591 1.87647 N 0.685 1.35 0.5 3.624339 6 13.7 scope49a-f.dat/41 85/15 0.273 0.590857 N 1.005 0.227 4.43 5.31746 38 39.9 scope48a-f.dat/42 85/15 0.439 0.950133 N 1.005 0.261 3.85 5.31746 85 50.5 scope47a-f.dat/43 85/15 0.53 1.14785 N 1.005 0.271 3.71 5.31746 108 55.2 scope45.mac/48 85/15 0.689 1.495675 N 1.002 0.229 4.37 5.301587 110.3 36.7 scope40.mac/49 85/15 0.548 1.188407 N 1.003 0.228 4.4 5.306878 157.1 65.4 scope36.mac/50 85/15 0.676 1.480755 N 0.993 0.3 3.31 5.253968 124.6 55.3 scope32.dat/51 70/30 2.22 1.678277 N 3.003 4.09 0.734 16.23243 34.76 64.0 npct31.dat/54 50/50 2.45 1.841726 N 3.02 9.78 0.31 16.32432 1.67 6.7 idmcap4.dat/56 60/40 1.801 1.817176 N 2.25 14.9 0.151 12.16216 0.81 6.7 npct36.dat.scope36.da 85/15 0.676 1.480755 N 0.993 0.313 3.173 5.253968 219 101.4 scoe37.dat/59 83/17 0.665 1.442137 N 1.003 0.301 3.332 5.306878 222 100.5 npct33.dat/60 85/15 1.625 1.557088 N 2.27 1.44 1.57 12.01058 101.8 90.2 scope29.mac/61 77.5/22.5 1.597 1.543861 N 2.25 1.886 1.19 11.90476 33.2 39.2 plastic1.dat/62 80/20 0.326 1.434592 N 1.13 1.26 0.897 9.04 21.81 84.3 npct28.mac/75 85/15 1.528 1.455805 N 2.283 1.23 1.9 12.07937 73.3 57.7 npct27.mac/76 80/20 1.555 1.478938 N 2.287 1.37 1.67 12.10053 48.8 43.0 scope26.mac/77 70/30 1.627 1.550807 N 2.282 2.141 1.07 12.07407 28.2 37.1 scope25.mac/78 70/30 1.659 1.577161 N 2.288 2.473 0.925 12.10582 17.6 26.2 scope 19. mac/79 70/30 1.519 1.428459 N 2.313 1.977 1.17 12.2381 34.3 44.6 npct18.mac/80 60/40 1.411 1.311586 N 2.34 5.101 0.46 12.38095 7.2 26.0 npct21.mac/81 50/50 1.77 1.659476 N 2.32 9.219 0.252 12.27513 1.8 9.4 npct24.dat/82 70/30 0.743 1.594373 N 1.003 2.4 0.42 5.278947 10.65 34.4 npct23.dat/83 70/30 0.754 1.617978 N 1.003 2.22 0.45 5.278947 10.85 31.9 npct20.dat/87 75/25 1.645 1.52609 N 2.32 2.544 0.912 12.21053 36.67 56.7