WO2001038170A1 - Fusee monoetage volant jusqu'a l'orbite - Google Patents

Fusee monoetage volant jusqu'a l'orbite Download PDF

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Publication number
WO2001038170A1
WO2001038170A1 PCT/US2000/005126 US0005126W WO0138170A1 WO 2001038170 A1 WO2001038170 A1 WO 2001038170A1 US 0005126 W US0005126 W US 0005126W WO 0138170 A1 WO0138170 A1 WO 0138170A1
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WIPO (PCT)
Prior art keywords
propellant
matrix
orbit
vehicle
solid
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Application number
PCT/US2000/005126
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English (en)
Inventor
Joe A. Martin
Larry H. Welch
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Technanogy, Llc
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Publication date
Application filed by Technanogy, Llc filed Critical Technanogy, Llc
Priority to AU36103/00A priority Critical patent/AU3610300A/en
Publication of WO2001038170A1 publication Critical patent/WO2001038170A1/fr

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/14—Space shuttles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/002—Launch systems
    • B64G1/006—Reusable launch rockets or boosters
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • B64G1/403—Solid propellant rocket engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/10—Shape or structure of solid propellant charges
    • F02K9/22—Shape or structure of solid propellant charges of the front-burning type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/26—Burning control
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/28—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants having two or more propellant charges with the propulsion gases exhausting through a common nozzle

Definitions

  • the present invention relates to rockets using solid propellant.
  • the invention relates to the use of solid propellant rocket motors to deliver a payload to earth orbit using a single engine stage.
  • 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 pa ⁇ loads into orbit around the Earth.
  • the launching system 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.
  • 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).
  • SSTO single stage to-orbit vehicle
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • flexibility in the structural design of the rocket will enable a single stage, solid fueled motor to accelerate a payload into Earth orbit.
  • 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.
  • 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.
  • 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.
  • 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.
  • the propellant cartridges may be constructed so that they contain only consumable materials and are completely consumed during their use.
  • 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.
  • FIGURE 1 shows a schematic view of a SSTO vehicle making use of a solid fueled rocket motor.
  • 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 (1 10) 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).
  • 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
  • Intimately mixed 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 solvatmg or dispersing the oxidizer in a solvent.
  • 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.
  • NRC-4 the metallic fuel is aluminum particles having an average diameter of about 40 nm, and the oxidizer is ammonium perchlorate (AP)
  • AP ammonium perchlorate
  • the aluminum and AP components of NRC-4 are present in stoichiomet ⁇ c 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.
  • the quantities of ammonium perchlorate and nanoaluminum were selected so as to yield a stoichiomet ⁇ c ratio of the ammonium perchlorate to the u ⁇ oxidized 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 Leyboid-Hereus TRIVAC D30A roughing/backing pump, and a 16- ⁇ nch diameter x 18- ⁇ nch 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.
  • 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.
  • reaction rates such as the burn rate of a particulate propellant mixture
  • reactant diffusion distance corresponds to particle size in particulate materials
  • 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.
  • the passivation layer When changing particle size, one must take the passivation layer into account in order to maintain the correct stoichiometry
  • the Al 2 0 3 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.
  • 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
  • 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.
  • 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 1 % by weight of the slower burning propellant component comprising hydrox ⁇ -terminated pol ⁇ butadiene (HTPB) and AP in stoichiometric quantities.
  • HTPB hydrox ⁇ -terminated pol ⁇ butadiene
  • any fuel/oxidizer propellant may be used, and mixed propellents may contain more than two propellant components.
  • 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:
  • R is the desired burn rate
  • m s is the mass of the slower burning propellant component
  • m f is the mass of the faster burning propellant component
  • R s is the burn rate of the slower burning propellant component
  • R f is the burn rate of the faster burning propellant component.
  • HTPB/AP is used as the slower burning propellant component due to its low cost, availability, and well-understood properties.
  • 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 HTPB/AP propellant has a higher net effect than each gram of a slower burning propellant component having a burn rate faster than HTPB/AP, as can be demonstrated using Equation 2.
  • 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.
  • 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
  • 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 solution/suspension, 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 miscibi ty 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
  • 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.
  • 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-m ⁇ crometer 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.
  • R h C P n , where R b is the burn rate, C is a constant, P is pressure, and n is the pressure exponent.
  • 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
  • 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 complete.
  • 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 elevated pressures that can be used to calculate the pressure exponent for the propellant.
  • a non-reactive gas e.g., dry nitrogen
  • Example 3 Burn Rate Testing and Pressure Exponent Determination of Propellant Mixture
  • the pressure vessel contained a pressure transducer (Endevco, 500 psig) and two electrical connectors to which a hot wire igniter (mchrome wire, 3 inches long by 0.005 inches in diameter) was attached.
  • 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)
  • 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-pressunzing 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.
  • 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 (Isp 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 and/or 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.
  • 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.
  • 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 Al 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
  • formulation 3 An additional factor which may be at work is the difference in the particle sizes.
  • the AP particles are, on the average, about 6-7 times larger than the Al particles.
  • formulation 5 the particles of Al 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.
  • 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.
  • the preferred propellants have very high energy, power, and burn rate as compared to propellants comprising more standard-like particle mixes.
  • mixed propellants comprising two components (i.e. propellants, fuel/oxidizer 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.
  • 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 3/4 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.
  • the propellant matrix (70) 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 bum-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.
  • 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.
  • 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
  • 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.
  • 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 structurally, the case need not provide as much structural support to the propellant as would be required in a CP design.
  • 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 09/447,758, filed on the same date as the present application, the entirety of which is hereby incorporated by reference.
  • 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.
  • 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.
  • the thrust profile is varied by varying the burn rate of the propellant matrix itself.
  • 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.
  • the thrust profile is varied by varying the size of the burn-front at different times during the burn.
  • 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 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
  • 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
  • propellant cartridges (250) Prior to launch, propellant cartridges (250) are mounted within each receptacle (240) of the transport vehicle
  • the payload is mounted in the vehicle, and then the vehicle is prepared for launch.
  • each propellant cartridge is ignited, and proceeds to burn in the manner discussed above.
  • 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.
  • 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
  • the transportation vehicle (210) 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 replaced 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.
  • the solid fueled rocket motor 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 launches.
  • FIGURE 4 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.
  • the propellant matrix (320) 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.
  • 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.
  • 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.
  • the propellant matrix (320) will be consumed. When the propellant is fully consumed, the cartridge will produce no more thrust and is completely expended.
  • 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 installed into the transport vehicle, and the system is prepared for another launch.
  • 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.
  • FIGURE 5 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
  • 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.
  • 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 nozzle on the cartridge, but may make use of a nozzle which is integral to the transport vehicle
  • propellant cartridges may be designed which incorporate some aspects of the self contained cartridge, and some of the fully consumable cartridge
  • a fully consumable propellant cartridge may be used in a transport vehicle which does not provide its own nozzle.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Remote Sensing (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Testing Of Engines (AREA)

Abstract

L'invention concerne un moteur de fusée à poudre capable de lancer un véhicule dans l'orbite terrestre au moyen d'un seul étage de moteurs n'utilisant qu'une seule matrice de combustible à taux de combustion élevé. Cette matrice de combustible est conçue pour se consumer d'une extrémité à l'autre, au lieu de se consumer à partir du centre. En fonctionnant avec cette configuration de combustion par les extrémités, ce moteur peut être configuré de manière qu'il produise divers niveaux de poussées pendant son fonctionnement. Il est ainsi possible d'utiliser le même moteur pour toutes les phases du lancement sur orbite.
PCT/US2000/005126 1999-11-23 2000-02-29 Fusee monoetage volant jusqu'a l'orbite WO2001038170A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
AU36103/00A AU3610300A (en) 1999-11-23 2000-02-29 Single-stage-to-orbit rocket

Applications Claiming Priority (2)

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US44792699A 1999-11-23 1999-11-23
US09/447,926 1999-11-23

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WO2001038170A1 true WO2001038170A1 (fr) 2001-05-31

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PCT/US2000/005126 WO2001038170A1 (fr) 1999-11-23 2000-02-29 Fusee monoetage volant jusqu'a l'orbite

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WO (1) WO2001038170A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2936570A1 (fr) * 2008-09-26 2010-04-02 Thales Sa Propulseur a combustible nano-energetique
FR3033770A1 (fr) * 2015-03-20 2016-09-23 Centre Nat D'etudes Spatiales (Cnes) Zone de lancement d'un engin aerospatial a propulsion a base d'un propergol solide et procede de protection associe
CN112661214A (zh) * 2020-09-30 2021-04-16 北京空间飞行器总体设计部 一种应对过载及背压的水升华器供水控制方法

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3165060A (en) * 1960-10-31 1965-01-12 James V Braun Multiple propellent grain for rocket motors
FR1510159A (fr) * 1966-12-09 1968-01-19 France Etat Nouveau bloc de propergol à combustion mixte
GB1266266A (fr) * 1969-08-27 1972-03-08
US3780968A (en) * 1967-01-03 1973-12-25 A Parilla Solid propellant rocket engine control and missile configurations
EP0335677A2 (fr) * 1988-03-30 1989-10-04 Satellite Network Tours, Inc. Lanceur et méthode pour lancer des objets dans l'espace
DE4422196A1 (de) * 1994-06-24 1994-11-10 Thomas Dr Stinnesbeck Anordnung zur Nutzung von Metallhydriden in Hybridraketentriebwerken
US5386777A (en) * 1992-02-10 1995-02-07 Aero-Jet General Corporation Rocket motor construction from porous binder core
WO1999047418A1 (fr) * 1998-03-17 1999-09-23 Davis Hubert P Lanceur recuperable avec module de propulsion par fusee amovible

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3165060A (en) * 1960-10-31 1965-01-12 James V Braun Multiple propellent grain for rocket motors
FR1510159A (fr) * 1966-12-09 1968-01-19 France Etat Nouveau bloc de propergol à combustion mixte
US3780968A (en) * 1967-01-03 1973-12-25 A Parilla Solid propellant rocket engine control and missile configurations
GB1266266A (fr) * 1969-08-27 1972-03-08
EP0335677A2 (fr) * 1988-03-30 1989-10-04 Satellite Network Tours, Inc. Lanceur et méthode pour lancer des objets dans l'espace
US5386777A (en) * 1992-02-10 1995-02-07 Aero-Jet General Corporation Rocket motor construction from porous binder core
DE4422196A1 (de) * 1994-06-24 1994-11-10 Thomas Dr Stinnesbeck Anordnung zur Nutzung von Metallhydriden in Hybridraketentriebwerken
WO1999047418A1 (fr) * 1998-03-17 1999-09-23 Davis Hubert P Lanceur recuperable avec module de propulsion par fusee amovible

Non-Patent Citations (4)

* Cited by examiner, † Cited by third party
Title
BEKEY I: "SSTO ROCKETS: A PRACTICAL POSSIBILITY", AEROSPACE AMERICA,US,AMERICAN INSTITUTE OF AERONAUTICS & ASTRONAUTICS. NEW YORK, vol. 32, no. 7, 1 July 1994 (1994-07-01), pages 32 - 37, XP000466206, ISSN: 0740-722X *
PELEGRIN M, HOLLISTER W M ET AL: "Concise Encyclopedia of Aeronautics & Space Systems", 1993, PERGAMON PRESS, XP002143639 *
RYAN R S ET AL: "FUNDAMENTALS AND ISSUES IN LAUNCH VEHICLE DESIGN", JOURNAL OF SPACECRAFT AND ROCKETS,US,AMERICAN INSTITUTE OF AERONAUTICS AND ASTRONAUTICS. NEW YORK, vol. 34, no. 2, 1 March 1997 (1997-03-01), pages 192 - 198, XP000693637, ISSN: 0022-4650 *
STANLEY D O ET AL: "PROPULSION REQUIREMENTS FOR REUSABLE SINGLE-STAGE-TO-ORBIT ROCKET VEHICLES", JOURNAL OF SPACECRAFT AND ROCKETS,US,AMERICAN INSTITUTE OF AERONAUTICS AND ASTRONAUTICS. NEW YORK, vol. 31, no. 3, 1 May 1994 (1994-05-01), pages 414 - 420, XP000456225, ISSN: 0022-4650 *

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2936570A1 (fr) * 2008-09-26 2010-04-02 Thales Sa Propulseur a combustible nano-energetique
FR3033770A1 (fr) * 2015-03-20 2016-09-23 Centre Nat D'etudes Spatiales (Cnes) Zone de lancement d'un engin aerospatial a propulsion a base d'un propergol solide et procede de protection associe
CN112661214A (zh) * 2020-09-30 2021-04-16 北京空间飞行器总体设计部 一种应对过载及背压的水升华器供水控制方法
CN112661214B (zh) * 2020-09-30 2022-08-05 北京空间飞行器总体设计部 一种应对过载及背压的水升华器供水控制方法

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