Notice of Pre-AIA or AIA Status
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
DETAILED ACTION
This Final Office action is based on the 19/013794 application originally filed January 08, 2025.
Claims 1-33, filed May 01, 2026, are pending and have been fully considered. Claims 17-33 are withdrawn from consideration due to being drawn to a nonelected invention.
Claim Rejections - 35 USC § 103
In the event the determination of the status of the application as subject to AIA 35 U.S.C. 102 and 103 (or as subject to pre-AIA 35 U.S.C. 102 and 103) is incorrect, any correction of the statutory basis (i.e., changing from AIA to pre-AIA ) for the rejection will not be considered a new ground of rejection if the prior art relied upon, and the rationale supporting the rejection, would be the same under either status.
The following is a quotation of 35 U.S.C. 103 which forms the basis for all obviousness rejections set forth in this Office action:
A patent for a claimed invention may not be obtained, notwithstanding that the claimed invention is not identically disclosed as set forth in section 102, if the differences between the claimed invention and the prior art are such that the claimed invention as a whole would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to which the claimed invention pertains. Patentability shall not be negated by the manner in which the invention was made.
The factual inquiries for establishing a background for determining obviousness under 35 U.S.C. 103 are summarized as follows:
1. Determining the scope and contents of the prior art.
2. Ascertaining the differences between the prior art and the claims at issue.
3. Resolving the level of ordinary skill in the pertinent art.
4. Considering objective evidence present in the application indicating obviousness or nonobviousness.
This application currently names joint inventors. In considering patentability of the claims the examiner presumes that the subject matter of the various claims was commonly owned as of the effective filing date of the claimed invention(s) absent any evidence to the contrary. Applicant is advised of the obligation under 37 CFR 1.56 to point out the inventor and effective filing dates of each claim that was not commonly owned as of the effective filing date of the later invention in order for the examiner to consider the applicability of 35 U.S.C. 102(b)(2)(C) for any potential 35 U.S.C. 102(a)(2) prior art against the later invention.
Claim(s) 1-16 is/are rejected under 35 U.S.C. 103 as being unpatentable over Hundley et al. (US 2024/0117782) in view of Elzein (US 2020/0240363).
Regarding Claims 1-16
Hundley discloses in the abstract, a multi-grained fuel grain for a rocket is disclosed, the method comprising the first propellant comprising a multi-grained fuel grain, the multi-grained fuel grain forming the at least one void, the at least one void facilitating variation in internal ballistics, forming sensors, said sensors permitting continuous monitoring and continuous modification such that a user controls the ballistics profile of a rocket motor, forming an electrically-controlled second propellant in contact with and operatively coupled to the sensors.
Hundley discloses in Figure 1A and 2A, depicts a fuel grain section with a generally cylindrical shape including a center chamber and a detailed fuel grain section with a generally cylindrical shape including the shell, center chamber, and annular sections and other elements.
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Hundley discloses in paragraph 0031, a method of making a fuel grain such as a multi-grained fuel grain for a rocket, the method requiring in an additive manufacturing process. This process uses at least one nozzle to form a multi-grained fuel grain that is composed of at least one propellant and a void that permits a user to program internal ballistics in a stepwise and/or continuous fashion. Further, sensors may be formed during this process, the sensors permitting continuous monitoring and continuous modification during the additive manufacturing process such that a user may control the ballistics profile of the rocket motor. Sensors may permit, for example, a user to remotely control the thrust and burning rate of a rocket motor in both a stepwise and continuous fashion, and also to indicate suitability of use over time of extended storage, a “health monitoring” system built into the SRM at time of production that allows nondestructive evaluation at any time.
Hundley further discloses in paragraph 0033, sensors are embedded into the fuel grain in order to facilitate continuous monitoring of the additive manufacturing process, the fuel grain including a plurality of layers of fuel grain material.
Hundley discloses in paragraph 0055, successive fuel beads are applied continuously by at least one nozzle 103 for a predefined distance in the direction primarily parallel to the central axis. Further, the multi-grained fuel grains may be formed by printing successive layers of dissolvable core, the multi-grained fuel grain and the dissolvable core printed layer by layer. Dissolvable core components may comprise but are not limited to a variety of binders, adhesives, metallic binders, and the like. Further, the multi-grained fuel grain may be supported by a fuel grain support structure, the fuel grain support structure comprising solid propellant and having unique geometrically shaped bead layers. In addition, the multi-grained fuel grain may be supported by a fuel grain support structure, the fuel grain support structure comprising a solid catalyst, solid fuels, and the like. Notably, the multi-grained fuel grain have unique geometrically shaped bead layers that facilitate user control of rocket ballistics.
Hundley discloses in paragraph 0058, sensors are embedded into the fuel grain in order to facilitate continuous monitoring of the additive manufacturing process, the fuel grain including a plurality of layers of fuel grain material, each layer including a plurality of concentric circular structures of different diameter. Further, forming the sensors includes extruding electrically-conductive material, said sensors including a conductive polymer material; and wherein forming the electrically-controlled second propellant material includes extruding the electrically-controlled solid propellant material. Further, in some embodiments the electrically-conductive material is an electrically-conductive fill in a polymer.
Hundley discloses in paragraph 0036, contemplates rocket combustion components such as an oxidizer, a rocket fuel, and a binder. Further, rocket fuel may include metal-based or non-metal based fuel. In addition, the center chamber may include unique patterns designed to induce oxidizer vortex flux persisting throughout the rocket's operation.
Hundley further discloses in paragraph 0072, a first propellant has a slower burn rate than a second or third propellant. In other embodiments, the propellant material may be any of a variety of materials that burn to produce pressurized gases, and that is ignited by using the sensors in the preferred embodiment. In some embodiments, rocket combustion components may comprise an oxidizer, a rocket fuel, and a binder. In some embodiments, the oxidizer may be a liquid- or solid-based perchlorate oxidizer that includes aluminum perchlorate, barium perchlorate, calcium perchlorate, lithium perchlorate, magnesium perchlorate, perchlorate acid, strontium perchlorate, and/or sodium perchlorate, singly or as eutectics or co-crystals with ingredients having positive oxygen balance. The rocket fuel may comprise a fuel grain (i.e., a solid propellant and solid fuel) and/or an incendiary fluid. Said fuel may be a metal-based or non-metal based fuel. Metal-based fuel may include, for example, tungsten, magnesium, copper oxide, copper, titanium, and/or aluminum. The binder may include casein, methyl cellulose, polyethylene oxide, polyvinyl acetate, and/or polyvinyl alcohol.
Hundley discloses in paragraph 0078, the rocket motor further comprises an oxidizer source. As described above, the oxidizer source is for flowing through the combustion port during motor operation. As described above, a shell is further utilized for housing at least the fuel grain and the oxidizer source.
It is to be noted, Hundley discloses a combustion port that has a shell that extends through the fuel grain assembly where the oxidizer flows through the port but fails to specifically define the combustion port is of hollow shape.
However, Hundley further discloses it would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to design a fuel grain assembly with the interior shape of the combustion port being a design choice depending upon the shape of the fuel grain assembly chamber. This is further taught in paragraph 0073 of Hundley, the rocket fuel comprises a fuel grain section contained in a rocket body, said fuel grain section having a generally cylindrical shape defining the center chamber, the fuel grain section comprising a solid propellant and a solid fuel about said solid propellant. In some embodiments, the center chamber has a substantially circular cross-section but in other embodiments a star shape or irregular shape is possible. It has been held to be within the general skill of a worker in the art to select a known material on the basis of its suitability for the intended use as a matter of obvious design choice. In re Leshin, 125 USPQ 416.
Hundley discloses in paragraph 0066, once the model is created and oriented, a preprocessing system then preferably divides the oriented fuel grain model into two-dimensional layers with defined footprint areas that increase with each subsequent two-dimensional layer until a width of the fuel grain model is achieved. The fuel grain production system receives the fuel grain model that is oriented in a direction of the central axis 115 and already divided into two-dimensional layers. The system then prints the fuel grain by applying at least one fuel layer in a direction primarily parallel to the central axis 115.
Hundley discloses in paragraph 0067, disclosed rocket utilizes a fuel grain that comprises a plurality of layers of fuel grain material, each layer including a plurality of concentric circular structures of different diameter fused together to form a central opening therein. Notably, in the preferred embodiment the fuel grain material comprises an ignitable substance. In other embodiments, the plurality of layers are stacked and joined securely to form a cylindrical fuel grain with the central opening of each one of the plurality of layers aligned to form a combustion unit extending axially through the fuel grain and bounded by a combustion surface. In some embodiments, the combustion unit comprises a circular shape, an oval shape, an elliptical shape, a cylindrical shape, a polygonal shape, or an irregular shape. Notably, the above-discussed central chamber is formed at least in part by said plurality of layers.
Hundley discloses in paragraph 0068, the fuel grain is configured to permit mixing of heterogenous materials to enhance thrust performance. In some embodiments, heterogenous materials comprise unfilled thermoplastics and/or filled thermoplastics, the unfilled thermoplastics and/or filled thermoplastics being filled with natural fibers, carbon fibers, minerals, and/or metal particles. Further, it is notable that heterogenous materials may comprise unfilled thermoplastics and/or filled thermoplastics with natural fibers, carbon fibers, minerals, metal particles, and heterogenous propellants.
Hundley discloses in paragraph 0080, each structural layer may be formed by deposition of viscous polymers, forming a plurality of solidified adjoining structural spheres of material. In some embodiments, the viscosity of said viscous polymers are optimized and controlled such that operational yield stress is enhanced relative to prior art systems. As for the adjoining sphere structures, the objective is to increase the surface area presented to the flame zone for combustion (in some embodiments, within the center chamber 102) in a manner that is persistent throughout the hybrid rocket engine operation.
It is to be noted, Hundley does not disclose the radial thickness of the solid propellant.
However, it is known in the art for a solid propellant of a fuel grain rocket engine to have a thickness of the radial direction that is less than the fuel grain material, as taught by Elzein.
Elzein discloses in paragraph 0117, a rocket engine encompassing a single solid propellant, being uniform and homogeneous, will burn very fast initially, then would start to burn slower and slower as the flame front moves along a regression direction. This may be explained by an increase of a volume of the combustion chamber with time as the solid propellant is gradually depleted. Therefore, with time, an increased volume of the combustion chamber may imply a lower concentration of the oxidizer in the combustion chamber, which may yield a lower regression rate. It may be possible to vary a mass flow rate of the oxidizer (e.g. to increase the mass flow rate) as the propellant is depleted to limit a decrease in the concentration of the oxidizer.
Elzein further discloses in paragraph 0118, a regression direction is a direction along which the solid propellant is being depleted. In other words, the regression direction represents a direction along which a dimension of the solid propellant decreases during burn. For instance, if the solid propellant is in the form of a tube defining a central passage, the regression direction may be a radial direction as a radial thickness of the tube decreases as the propellant is burned away during the combustion process. In such a case, a dimension of the central passage may increase with time during the combustion process. As another example, if the solid propellant is in the form of a solid cylinder having an axial circular end face exposed to the oxidizer, the regression direction may be an axial direction as a length of the solid propellant decreases during the combustion process. Other configurations are contemplated.
Elzein discloses in paragraph 0119, the combustion chamber 1304 contains a plurality of fuels each having a solid propellant. In the embodiment shown, the combustion chamber 1304 contains three fuels, namely a first, a second, and a third fuel 1318 a, 1318 b, 1318 b that are axially stacked one next to the other relative to the longitudinal axis L of the engine 1300. These three fuels 1318 a, 1318 b, 1318 c are, in the present embodiment, cylinders or disks. Each of these three fuels 1318 a, 1318 b, 1318 c is made of a solid propellant differing from that of the others by their regression rates. This may be achieved, for instance, by having the three fuels differing by one or more rheological properties. It would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art for the solid propellant of Hundley to have a thickness of the radial direction that is less than the fuel grain material, as taught by Elzein in order to improve the burn rate and decreases the radial thickness of the propellant as it is burned away during the combustion process.
Response to Arguments
Applicant’s arguments directed to the newly submitted claimed amendments, filed May 01, 2026, with respect to the rejection(s) of claim(s) 1-16 under Hundley et al. (US 2024/0117782) have been fully considered and are persuasive. Therefore, the rejection has been withdrawn. However, upon further consideration, a new ground(s) of rejection is made in view of Hundley et al. (US 2024/0117782) in view of Elzein (US 2020/0240363).
Conclusion
The prior art made of record and not relied upon is considered pertinent to applicant's disclosure.
Danforth et al. (US 2017/0253537) discloses in the abstract, a propellant material element, such as an electrically-operated propellant material, includes extruding a propellant material through a heated nozzle. The nozzle may be heated to a temperature that is above the boiling point of a solvent that is part of the propellant material, yet is below a decomposition temperature of the propellant material.
Danforth et al. (US 2010/0307131) discloses in the abstract, a solid fuel rocket motor, a castellated propellant cartridge and a method of controlling a pressure differential in a cartridge-loaded rocket motor are disclosed. The rocket motor may include a housing having an inside surface, a plurality of propellant cartridges disposed within the housing, an igniter disposed to ignite propellant material within the propellant cartridges, and a nozzle disposed to exhaust combustion gases out of the housing.
Schumacher et al. (US 5,714,711) discloses in the abstract, a propellant grain composition for use in solid fuel rocket engines comprises a first reactant component consisting of an oxidizer. A polymeric barrier coating encapsulates the first reactant component and a second reactant component comprising a reducer fuel react with the oxidizer is coated thereon, a final polymeric coating is placed over the entire propellant grain to yield a unitary metal fuel-oxidizer propellant grain structure for use as a solid rocket fuel.
Applicant's amendment necessitated the new ground(s) of rejection presented in this Office action. Accordingly, THIS ACTION IS MADE FINAL. See MPEP § 706.07(a). Applicant is reminded of the extension of time policy as set forth in 37 CFR 1.136(a).
A shortened statutory period for reply to this final action is set to expire THREE MONTHS from the mailing date of this action. In the event a first reply is filed within TWO MONTHS of the mailing date of this final action and the advisory action is not mailed until after the end of the THREE-MONTH shortened statutory period, then the shortened statutory period will expire on the date the advisory action is mailed, and any nonprovisional extension fee (37 CFR 1.17(a)) pursuant to 37 CFR 1.136(a) will be calculated from the mailing date of the advisory action. In no event, however, will the statutory period for reply expire later than SIX MONTHS from the mailing date of this final action.
Any inquiry concerning this communication or earlier communications from the examiner should be directed to LATOSHA D HINES whose telephone number is (571)270-5551. The examiner can normally be reached Monday thru Friday 9:00 AM - 6:00 PM.
Examiner interviews are available via telephone, in-person, and video conferencing using a USPTO supplied web-based collaboration tool. To schedule an interview, applicant is encouraged to use the USPTO Automated Interview Request (AIR) at http://www.uspto.gov/interviewpractice.
If attempts to reach the examiner by telephone are unsuccessful, the examiner’s supervisor, Prem Singh can be reached at 571-272-6381. The fax phone number for the organization where this application or proceeding is assigned is 571-273-8300.
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/Latosha Hines/Primary Examiner, Art Unit 1771
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