Classifications

• • 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
  • F02K9/30—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants with the propulsion gases exhausting through a plurality of nozzles
• • 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
  • F02K9/97—Rocket nozzles
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F42—AMMUNITION; BLASTING
  • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
  • F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
  • F42B15/10—Missiles having a trajectory only in the air
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
  • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
  • F05D2240/00—Components
  • F05D2240/40—Use of a multiplicity of similar components
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
  • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
  • F05D2300/00—Materials; Properties thereof
  • F05D2300/20—Oxide or non-oxide ceramics
  • F05D2300/21—Oxide ceramics
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
  • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
  • F05D2300/00—Materials; Properties thereof
  • F05D2300/20—Oxide or non-oxide ceramics
  • F05D2300/22—Non-oxide ceramics
  • F05D2300/224—Carbon, e.g. graphite
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
  • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
  • F05D2300/00—Materials; Properties thereof
  • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
  • F05D2300/603—Composites; e.g. fibre-reinforced
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
  • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
  • F05D2300/00—Materials; Properties thereof
  • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
  • F05D2300/614—Fibres or filaments

Definitions

• the field of the subject invention are jet propulsion systems, and more particularly the invention pertains to the use of a multi-nozzle grid to direct transonic and supersonic flows in rocket motors for use in propulsion.
• Jet propulsion uses the momentum of ejected matter to propel a vehicle or device, the ejected matter usually being predominately a gas. Rocket motors are one of the most common applications of jet propulsion. Rocket motors propel vehicles or devices called rockets or missiles .
• jet propulsion engines that enhance the flight characteristics of the vehicle. Stability in flight is aided by having the center of gravity ahead of the center of aerodynamic pressure. Otherwise, alternative means of adding stability, such as tail fins, must become necessary to achieve a desired level of stability.
• Rocket motors most commonly burn solid (“solid rocket motor” or “SRM”)or liquid (“liquid rocket motor” or “LRM”) fuels contained in the rocket to produce very high temperature gasses which are ejected from the rocket engine at several times the speed of sound.
• SRM solid rocket motor
• LRM liquid rocket motor
• the length of operation for an engine is also important .
• a rocket engine needs to operate long enough to accomplish an objective, such as delivery of a payload. Accordingly, jet propulsion engines that offer longer operational periods are desired. Longer operation can occur through longer survival of the rocket engine under the stresses of extremely hot and high pressure fuel, or from a weight savings that allows a fixed amount of fuel to be used over a longer period of time.
• Nozzles generally can have a convergent section where the nozzle accepts gasses, a throat which is the most constricted part of the nozzle, and a divergent section where gasses are expanded prior to being expelled from the engine. Some nozzles may have only convergent or divergent sections, however such nozzles will not have practical uses for gasses expelled at supersonic speeds. In the case of gasses expelled at supersonic speeds, the selection of the parameters for the divergent nozzle are more important than the convergent nozzle.
• Other general consideration in jet propulsion engine design include allowing for the fact that discontinuities on the walls of the engine are likely to give rise to . energy losses from shock waves, so all nozzle sections should be well rounded. The exit portion of the divergent section usually has a sharp edge because a rounded edge would permit overexpansion and flow separation in the expelled gases.
• the use of multiple conventional nozzles in rocket design is known, but not favored by those of ordinary skill in the art.
• Multiple conventional nozzles have been used when the geometry (i .e. , length) or weight of a single conventional nozzle was prohibitive. While the concept has been generally limited to small tactical missiles in the western hemisphere, the use of multiple conventional nozzles was applied even for space exploration in the eastern hemisphere, especially in the Soviet Union. Yet, the use of multiple conventional nozzles is generally considered by those of ordinary skill in the art to be less efficient than using a single nozzle. Because increasing the number of nozzles is generally thought by those of ordinary skill in the art to increase inefficiencies, even when multiple conventional nozzles are used, the use is usually limited to four to six nozzles at most to minimize the generally perceived disadvantages of multiple nozzles.
• nozzlettes An early use of a plurality of small nozzles (“nozzlettes”) was applied in a supersonic wind tunnel in Germany during the 1930s to overcome length limitations [1] .
• the construction was to place a rectilinear grid. of orifices in a substantially rectangular wind tunnel.
• the use of multiple nozzlettes achieved a length savings, but forced the designers to use a settling chamber with the length of an equivalent single nozzle because the Germans did not have a knowledge of fluid dynamics that would permit them to control the scale and decay distance of turbulence by the selection of the number and size of nozzlettes .
• Typical problems of conventional jet propulsion engine design include geometrical limits imposed on the nozzle length and/or diameter, the weight limitations of an efficiently designed ideal single nozzle, the requirement of the selection of heavy material for throat design and the deleterious aerodynamic effects of an aft shifting of center of gravity on the aerodynamic stability of the air vehicle.
• a design procedure that allows dramatically increased performance, adherence to theoretically superior nozzle geometries, and reducing weight while also cutting the cost and time to manufacture jet propulsion engines would meet needs not met adequately by current technology.
• the present invention relates to an improved nozzle system for use in propulsion.
• One aspect of the present invention relates to methods of designing multiple nozzlette plates for use in propulsion.
• Another aspect of the present device is a multi-nozzle grid for use in jet propulsion, whether rocket, jet turbine, or other, that provides structural integrity to a jet propulsion device while aiding the management of drag from the gas ejected to propel a device.
• One aspect of the present invention is a jet propulsion outlet device comprising a grid plate having a plurality of densely clustered nozzlettes, the nozzlettes of the grid plate being configured to operably couple to a pressurized gas source to efficiently expand the pressurized gas.
• a jet propulsion outlet device with a plate that is made from a material from the group consisting of glass reinforced phenolic composites, graphite reinforce phenolic composites, short strand reinforced phenolic composites, fiber reinforced ceramic matrix composite, and ceramic composites .
• the jet propulsion outlet device has nozzlettes that are disposed in a pattern having a port to nozzlette ratio of greater than one .
• the jet propulsion outlet device wherein the nozzlettes are made of a material that will remain substantially intact after having a gas stream having a pressure of 14,000 psi and a temperature of 2000 °C for 120 seconds passed through the nozzlettes .
• the jet propulsion outlet device has at least one centrally disposed nozzlette surrounded by a plurality of peripheral nozzlettes, each of the plurality of peripheral nozzlettes abutting at least one central nozzlette and at least two other peripheral nozzlettes.
• the jet propulsion outlet device has nozzlettes that are disposed in a pattern such that when a pressurized gas is passed through the nozzlettes, the pattern is substantially free of stagnation zones and the pressurized gas is not subjected to flow turning.
• the jet propulsion outlet device has nozzlettes where the convergent portion of the nozzlettes converges at an angle of less than 48°, and the divergent portion of the nozzlettes diverges at an angle of less than 30°.
• Another aspect of the present invention relates to methods of designing a nozzlette grid for channeling a gas comprising the steps of: providing design parameters; determining a required plate thickness based on the design parameters ,- determining a geometry of an equivalent single nozzle; defining geometric pattern to pack the nozzlettes in a tight arrangement; and selecting a number of nozzlettes; wherein the design for the nozzle grid defines a plate having the required plate thickness having the plurality of nozzlettes with the geometry of the equivalent single nozzle disposed in the geometric pattern.
• the design parameters include parameters related to mechanical and thermal stresses associated with the application of a gas to the nozzlette grid and the materials properties of a material.
• the geometric pattern is such that when a pressurized gas is passed through the nozzlettes, the defined plate is substantially free of stagnation zones and the gas is not subjected to flow turning.
• Another aspect of the present invention is related to missiles having improved aerodynamic stability having a payload, a propellant, and an engine comprising a plate having a plurality of nozzlettes disposed in a pattern that reduces stagnation zones in the engine.
• the missile also has a center of gravity of the payload, engine, and unexpelled propellant, where the center of gravity being spaced from the engine.
• the missile also has a center of aerodynamic pressure, the center of aerodynamic pressure being located closer to the engine than the center of gravity.
• a preferred embodiment of the present invention relates to a missile in which the engine has a center of gravity that is further forward than that of an equivalent single nozzle engine made from the same material .
• a missile comprising a payload and a propellant, the propellant being capable of being a pressurized gas.
• the missile also has an engine comprising a grid plate having a plurality of densely clustered nozzlettes, the nozzlettes of the grid plate being configured to operably couple to a pressurized gas source to efficiently expand the pressurized gas.
• a missile has a motor having a mass less than that of an equivalent single nozzle engine made from the same material .
• the missile has a plate that is made from a material from the group consisting of glass reinforced phenolic composites, graphite reinforce phenolic composites, short strand reinforced phenolic composites, fiber reinforced ceramic matrix composite, and ceramic composites .
• the nozzlettes of the missile are made of a material that will remain substantially intact after having a pressurized gas having a pressure of 14,000 psi and a temperature of 2000 °C for 120 seconds passed through the nozzlettes .
• the missile has at least one centrally disposed nozzlette surrounded by a plurality of peripheral nozzlettes, each of the plurality of peripheral nozzlettes abutting at least one central nozzlette and at least two other peripheral nozzlettes.
• the missile has nozzlettes that are disposed in a pattern such that when a pressurized gas is passed through the nozzlettes, the pattern is substantially free of stagnation zones and the pressurized gas is not subjected to flow turning.
• the convergent portion of the nozzlettes converges at an angle of less than 48°, and the divergent portion of the nozzlettes diverges at an angle of less than 30°.
• the present invention has several benefits and advantages .
• the methods and apparatus of the present invention can be used to reduce the length and weight of gas inlet and outlet management devices for jet propulsion. This in turn can provide jet propulsion engines, including rockets and turbines, having superior specific impulse characteristics.
• the present invention can also be used to control the scale and decay-distance of turbulence in jet propulsion.
• the present invention can provide missiles having improved aerodynamic stability.
• the present invention is capable of sustaining reasonable burn times for rocket, propulsion.
• the present invention is capable of provide superior structural strength in jet propulsion application, while providing other benefits such as preventing the intake of foreign objects into jet turbine engines .
• Fig. 1 is a bottom perspective view of a missile embodying the present invention
• Fig. 2 is a bottom perspective view of a rocket motor embodying the present invention
• Figs. 3A-B are sectional views of flat (3A) and convex (3B) grid plates embodying the present invention
• Fig. 4 is a perspective view of a single nozzle illustrating varying efficiency levels of single nozzle of different lengths
• Fig. 5 is a schematic of a single nozzle illustrating many of the parameters that are used to define such a nozzle
• Fig. 6 is a diagram of the arrangement of circles within a circle to provide a centrally disposed pattern of nozzles
• Fig. 7 is an illustration of prior art conventional multiple nozzle arrangements
• Fig. 8 is an illustration of nozzlette arrangements of the present invention.
• Figs. 9 A-B are above and side schematics of a tested one nozzle configuration
• Figs. 10 A-B are above and side schematics of a tested seven nozzlette configuration.
• FIGs. 11 A-B are above and side schematics of a tested nineteen nozzlette configuration.
• Fig. 1 shows an embodiment of the present invention in the form of a missile 10.
• the missile 10 of Fig. 1 is a single-stage rocket having a tail section 12.
• the tail section has a source of pressurized gas 14 or other expelled matter operably connected to a multi-nozzle grid plate (referred to as "multi-nozzle grid” or "MNG") 16 through which the expelled matter is sent.
• MNG multi-nozzle grid
• this operable connection is a chamber 18 located between the source of pressurized gas 14 and the MNG 16.
• the missile can optionally have aerodynamic elements such as fins 20 to add stability or steering capabilities in flight.
• a missile can comprise one or more stages each having a tail section 12, and each tail section having a pressurized gas source 14 and a nozzle grid 16.
• Fig. 2 is a bottom perspective of a MNG 16 engine 21 (or motor, or more generally than rockets, outlet device) of the present invention having 201 nozzlettes 22 and defining a chamber 18.
• the MNG 16 has a thickness 24.
• Fig 3A shows a flat partial cross-section of the nozzle grid 16 of Figs 1 & 2, each nozzlettes 22 can have a convergent section 26, a throat 28 and a divergent section 30.
• the source of pressurized gas 14 directs pressurized gas through the nozzlettes 22 to propel the missile 10.
• Fig. 3B shows a MNG 16 that rather than being flat has a convex geometry.
• a MNG 16 can be specified through a design procedure that uses :
• the design procedure described above can specify a MNG plate 16 that is thinner and lighter than a single nozzle.
• the length saving is in proportion to the square root of the number of the nozzlettes 22 in the MNG 16 (i.e., a MNG with 100 nozzlettes is about 10 times thinner than an equivalent single nozzle) .
• the multi nozzle grid 16 reduces energy losses to flow speed losses and heat transfer losses.
• the present invention's placement of the nozzlettes 22 reduces or eliminates stagnation zones.
• the MNG 16 accomplishes the reduction of stagnation zones while providing a structural element that can provide structural stability to a device such as a missile 10, turbine, or other kind of jet engine. It is also thought that the a MNG 16 of the present invention reduces flow turning at the outlet of the engine, and thereby avoids losses owing to drag and heat transfer inherent in turning a gas stream.
• missiles 10 that could improve performance and reduce production cost by using a MNG 16 configuration instead of rocket nozzles of the prior art.
• one type of missile 10, interceptors can improve their terminal velocity or reduce mass and size for the same performance.
• an interceptor missile that achieves a high burnout velocity, if designed with a MNG 16 might be small enough to fit into existing platform instead of going to a larger platform.
• the example is not limited, and the MNG 16 of the present invention can be used for both tactical and ballistic missiles 10.
• the MNG 16 design procedure has been used successfully in tactical missiles 10 using both stainless steel and short strand glass reinforced phenolic composite.
• the recent arrival of heat-resistant materials (for hypersonic flight of scramjet engines, turbine and wheel brake pads of passenger airplanes) provides an inventory of heat- and erosion-resistant materials that can operate much longer than practical application, such as, but not limited to, missile defense interceptors require .
• Figs. 2-3 presented embodiments of the present invention as advanced rocket motors with an MNG 16 configuration.
• Fig. 4 shows a conventional single-nozzle rocket motor 32 having three possible different lengths for the single nozzle.
• a practical single nozzle 34 which signifies a conventional engineering choice, is seen as the shortest embodiment.
• An optimal single nozzle 36 which can be defined as being adapted for an anticipated expansion ratio where the exit pressure equals the ambient pressure, is longer.
• an equivalent single nozzle 38 that is proportionally sharing identical geometrical properties with each individual nozzlettes 22 of the MNG 16 is the longest.
• Fig. 2 While the advanced rocket motor 21 of Fig. 2 consists of a compact chamber 18 with a MNG plate 16 that is short, compared to the longer equivalent single nozzle 38 of Fig. 4, the details of Figs 3A-B reveal that the MNG 16 has many nozzlettes 22. These nozzlettes 22 can have the same scaled-geometry as that of the equivalent single nozzle 38 of Fig. 4.
• SRM solid fueled rocket motors
• application of the principles of the present invention is not limited to SRM design. And can be applied to other types of rocket propellants such as liquid fuel propellants as well as several applications in other engines, including, but not limited to, the jet turbines to be discussed.
• a conventional SRM with a practical single-nozzle 34 (i.e., one that considers mass and geometric limits) must be much shorter than that of the equivalent single nozzle 38 because of the expansion ratio limits. These limits are controlled by several factors, including, but not limited to, 1) missile diameter; 2) ambient pressure outside the rocket; and 3) the reduction in missile velocity due to the extra weight of an added portion of the nozzle [4.
• mass properties considerations that are generally very important to missile design, a lighter aft body improves aerodynamic static stability by moving the center of gravity forward. Alternatively, length saving obtained can provide improved performance by simply adding more propellant.
• the multi nozzle grid 16 design procedure includes a equivalent single-nozzle design 40 illustrated in Fig. 5.
• the design also considers the thrust coefficient, Cf, which is an important element in ideal nozzle design, that relates the predicted performance and requirements to nozzle geometry.
• Cf the thrust coefficient
• handbooks of solid rocket design [4,5,6,7,8] detail ways to design nozzles such that the thrust coefficient is optimal.
• the geometric design procedure of the MNG procedure is also included.
• the second term is applicable in two cases: 1) P2>P3 for under-expanded nozzle, or 2) P2 ⁇ p 3 f° r over-expanded nozzle.
• Nozzles with exceptionally high expansion ratios are usually useful for exo-atmospheric applications.
• P3 0 (i.e., the vacuum of space) there is no limit of over expansion.
• P2 ⁇ p 3 the nozzle is not efficient because the flow separation due to negative pressure on the nozzle exit tips, reduces the effective " expansion ratio.
• the second term is then negative and the value of Cf diminishes.
• the over-expanded nozzle is wasteful and an under-expanded nozzle is more practical, not only because Cf cannot be reduced further by the P e -Po term, but also because the geometric area ratio of the exit to throat ( ⁇ ) limits. This is also true to jet turbine and some other non-rocketry applications of the present invention. This ratio ( ⁇ ) , which is limited by length constraints in conventional nozzle design, can be exploited using the MNG configuration.
• the design of the MNG can begin with a standard single-nozzle design as shown in Fig. 5.
• This equivalent single nozzle design 40 can conform to all the textbook design criteria for nozzles such as, but not limited to, those known to those of ordinary skill in the art [4,5,6,7,8,9].
• This step can also beneficially include calculations of burn surface and initial void-volume in the chamber.
• MNG design is especially sensitive to void volume changes due to its significant reduction in convergent nozzle volume. Void volume controls the initial pressure transient and can be easily obtained using ref. [9].
• Fig. 5 shows an equivalent single nozzle.
• the MNG procedure can describe this equivalent single nozzle according to the following equations:
• A is the throat area of the equivalent single nozzle
• Eq. (10) shows that the length saving of MNG configuration is proportional to the square root of the number of nozzlettes selected. For example, MNG with 196 nozzlettes will be about fourteen (14) times shorter than that of an "equivalent single nozzle.” For example, one MNG configuration was successfully tested used 201 nozzlettes. Eq. (10) then helps quantify the large value of length saving can be achieved by increasing the number of nozzlettes in the MNG configuration.
• the maximum number of nozzlettes can be determined by how many nozzlettes can be fit into this thickness. Following standard design procedure (i.e., safety factor, etc.), the maximum number of nozzlettes is determined thereafter [10] . Knowing now both L of an equivalent single nozzle and L MNG yields:
• Port area is defined by the cross-sectional area of hot gases and combustion particulate from the surface of the solid propellant or the liquid injectors of oxidizers and fuels towards the nozzle throat.
• the flow converges, unobstructed from rest in the far flowfield to sonic speed in the nozzle throat.
• the burn surface is changing and the reference area that defines the starting line progressively recedes away from the initial burn surface .
• the burn surface is not always limited to burn- back configuration (i.e., where the flow proceeds away from the burn surface that is perpendicular to the nozzle throat, in a straight line from the surface until it exit through the nozzle) • More often than not, the burn surface is parallel to the centerline.
• burn- back configuration i.e., where the flow proceeds away from the burn surface that is perpendicular to the nozzle throat, in a straight line from the surface until it exit through the nozzle
• the burn surface is parallel to the centerline.
• a tube geometry or a cluster of tubes where the burn surface is mostly occurring on the internal or external round surfaces. In this case, the flow is forced to turn in a right angle before being accelerated towards the nozzle.
• the port in this case does not match the burn surface.
• the burn surface is the tube internal surface plus the ring facing the nozzle (assuming the outer surface is bonded to the chamber pressure wall)
• the port is the cross section area of the flow exiting the tube on its converging way towards the nozzle throat.
• the port area still conforms to that definition (i.e., the combustion chamber internal cross section minus the obstruction area) .
• the port to nozzle ratio should go to infinity [7] .
• value close to one are most common.
• Local port to nozzle ratio in MNG geometry refers to the contribution of a single nozzlette. It is therefore easy to see that when the number of nozzlettes increases this ratio goes to infinity.
• Figs. 6 A-D show four exemplary arrangements of circular nozzlettes 22 within a circular nozzle grid plate 16. Of the four, Fig. 6A provides the most tightly packed grid having nineteen nozzlettes. The small circles represent the exit diameter 42 of each nozzlette 22 and can be calculated following the equations given above. These formulae are known for other purposes to those of ordinary skill in the art, for example, the formula to define Pattern A in Fig. 6 being defined in ref . [11] .
• the nozzlette 22 pattern comprises a core of centrally disposed nozzlettes 43 surrounded by one or more rings or layer of peripheral nozzlettes 45.
• the centrally disposed nozzlettes will have a high degree of symmetry to add stability to the in-flight stability of the rocket. More preferably, as shown in Figs. 6A-C, the nozzlettes are disposed in a hexagonal arrangement. As shown in Fig. 6A, there is but one central nozzlette, giving rise to an arrangement of a hexagon with three nozzlettes on a side. An alternative arrangement, more diamond-shaped than Fig. 6A, shown in Fig 6B is less symmetrical, more like a diamond shape, but still contemplated by the present invention The arrangement shown in Fig. 6C has a triangular arrangement of centrally disposed nozzlettes 43 that gives rise to a more generally triangular nozzlette pattern. The arrangement in Fig. 6D is more rectangular than Fig. 6A.
• the approach to form a densely clustered pattern of nozzlettes is to have as many nozzlettes packed substantially as tightly as practical.
• the centrally disposed nozzlettes can be arranged to touch in a touching or almost-touching formation, as seen in Figs. 6B-6D.
• a maximum number of peripheral nozzlettes can then be placed adjacent to and abutting the central nozzlettes to maximize the density of nozzlettes in a port area 47.
• the nozzlette pattern will substantially span the port area 47.
• Cf The calculations of Cf can follow the formulation detailed in reference [6] , section "Thrust and Thrust Coefficient," p. 58-63.
• Chamber pressure (P ⁇ ) is constant while the exit pressure is allowed to vary in order to generate a series of Cf's, ⁇ 's and F's.
• the mass flow rate, wdot (usually depicted in texts as a w with a dot over it and having often units of lb/sec) is a constraint based on density and bu n-time.
• a e /At determines which of the nozzle expansion ratios ( ⁇ ) is appropriate.
• the burn area to throat area ratio, K n is also calculated based on equation 11-13, p. 384 [6] . The calculations can be done by hand, or more conveniently using commercially available software such as Mathcad.
• Figs. 7 & 8 illustrates this point in relations to the MNG configurations.
• rocket motors 44 with multi nozzle arrangement away from the center i.e., from 18 nozzles 46 in a circle close to the circumference in the Russian made Katusha (Fig. 7C) to a four nozzles 46 in the MK 72 (Fig. 7A) and many other multi nozzle examples) suffer from losses due to the flow turn from the center to the orifices away of the centerline 48.
• Cluster design consists of separate rockets each having its own combustion chamber and nozzle/nozzles. This practice is probably as old as the first rockets that were produced for the Chinese Emperors millennia ago. When higher fire power or longer range was needed and the only available inventory was of smaller caliber, cluster was a quick fix that represented manufacturing compromise.
• cluster is a systematic 'packaging' of many small caliber rockets into a single unit.
• the systematic packaging of many small nozzlettes into a single unit of MNG is quite similar. Since the MNG shares a single combustion chamber, as opposed to the many small caliber combustion chambers of each individual rocket in the cluster, the structural mass saving of the MNG is readily recognized by those of ordinary skill in the art. Calculations show that the MNG with n nozzlettes is lighter than a cluster having the same n number of rockets, same material and overall similar diameter and thrust level .
• Composites and other materials with densities similar to that of the propellant are, in general, desirable choices for the rocket motor structure, not only because of the superior yield stress to density ratio composites display, but also because of the effect on the resulting overall mass reduction property of the missile.
• the present invention is not bound by any particular limit of yield stress to density ratio. It is rather limited by the suitability of the material, which in response to the intense heat can erode excessively and/or unacceptably (i.e. above 10%).
• yield stress to density ratio of composites is in the range of 30 to 5 million lbf/lbm are suitable for the present invention as compared to stainless steel which ranges from a million to 100,000 lbf in/lbm, and is not acceptable for all applications of the present invention.
• MNG technology is preferably made from composite materials. Instead of expensive machining, a matrix akin to mass-produced casting can result in a single part. For example, an MNG plate comprising the MNG and it associated case would drastically reducing production costs.
• the material is preferably a Glass or Graphite reinforced phenolic composite with or without multi-ply woven fabric inserts.
• vacuum plasma spray of thin layer of heavy ceramics or metals over the composite matrix can provide beneficial performance characteristics including much longer burn times. As shown in Table 1, below, materials containing or treated with niobium compounds, such as Columbium C103 can provide very long burn times.
• Transfer molding of short strand reinforced phenolic with a MNG plate thickness of only 1/4 inch has shown to safely last for 5 seconds.
• use of a 2- inch thick MNG plate can extend the operating time to over 10 seconds.
• transfer molding with multi-ply graphite woven fabric inserted in the throat area reduces nozzle erosion to 3%.
• ceramic inserts for every individual nozzlettes convergent cone can be placed above the Phenolic impregnated graphite woven fabric in the matrix, before the transfer molding (RTM) process begins.
• carbon-carbon matrix and ceramic-carbon (C/SiC) composite material [11,12,13] can be used to fabricate the whole MNG plate separately or as an integral part of the pressure chamber. Tests operated from 36 to 56 seconds exhibit acceptable results with some nozzle erosion [11, page 228] . Rocket motors for space exploration, which used columbium alloy C103 at a working pressure of 1800 psi and temperature of 2300° F, were reported to operate for over 900 seconds without apparent degradation [12] . More recent studies reported testing material at 1500 psi and an operating temperature of 3000° Kelvin show 21 seconds operation without erosion [13] .
• Ceramic compounds are silicon based and have exceptional endurance in high temperature applications. Many of the ceramics available for practice in the present invention were developed during the efforts to develop hypersonic flight worthy components during the last few decades, and the suitability of a compound for use in the present invention can be informed from the published literature concerning such development. Some of the ceramic compounds are enriched with carbon, zirconium and metals such as aluminum in order to enhance one property or another. Some ceramics' densities are somewhat higher than composites, but still much lower than that of a metal, leaving them still suitable for practice of the present invention.
• Heavier metal ceramic i.e., Rhenium, Tantalum carbide, Hafnium carbide, Hafnium diborate and Hafnium nitride
• Mold sintering production method for ceramics is another option for mass-producing nozzlettes' convergent inserts [11,12,13,15].
• multi nozzle grids of the present invention can be made by forming nozzlettes of a suitable material, and embedding them in a plate or assembling them in an array by methods known to those of ordinary skill in the art.
• such design must necessarily take into account the ability of the final product to withstand the stresses of the particular application. For example, with respect to rockets, the assembly must survive the heat and pressure of the propellant being expelled.
• Table 1 shows the relative durability of several different materials when exposed to solid rocket burn conditions .
• Suitable materials can include fiber reinforced ceramic matrix composite materials that can be obtained from Ceracom from the Ceramight ENVI fiber reinforced ceramic matrix composite model line. Such products can have 2-D or 3-D fiber weaves, and can be made from, but are not limited to, SiCf/SiC, SiCf/SiC+Si, Cf/SiC and matrices: HfC, HfN, TaC, B 4 CF. Other suitable materials available from Ceracom include ceramic composites sold under the CERAMIGHT brand. The CERAMIGHT materials can have bending strengths of more than 180 MPa at 20 °C, more than 140 MPa at 1500 °C, or more than 80 MPa at 2000 °C.
• Example 2 Further studies using the multi-nozzle grid continued to prove the advantages found in Example 1.
• the studies of Example 2 used a multi -nozzle grid having 201 nozzlettes formed from stainless steel as part of a solid propellant rocket engine.
• the nozzle length (14:1) and the nozzle weight (5:1) were drastically reduced as compared to practical single nozzles.
• the MNG tactical booster motor of Example 2 was operated at a pressure of 14,600 psi and is illustrated in Fig. 3.
• the MNG of Example 2 boosted a missile to a muzzle velocity that was more than 30% higher than a conventional configuration with a practical single nozzle as explained in conjunction with the description of Fig. 4.

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• Engineering & Computer Science (AREA)
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• General Engineering & Computer Science (AREA)
• Mechanical Engineering (AREA)
• Aviation & Aerospace Engineering (AREA)
• Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
• Nozzles (AREA)
• Testing Of Engines (AREA)

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