Office Action 1: Non-Final Rejection (2025-01-29)
Rejection types: 102 - Lack of Novelty

CLAIMS:
AMENDMENTS TO THE CLAIMS

Application No.: 18/749,428 Filing Date: June 20, 2024

1\. (Cancelled)
2\. (New) A solar thermal rocket engine, comprising:
a solar thermal thruster body;
one or more mirrored collectors configured to concentrate solar energy into an energy beam; and
an optical aperture configured to modulate an amount of the energy beam allowed to pass into the solar thermal thruster body.
3\. (New) A method of starting up a solar thermal rocket engine, comprising:
providing a gaseous propellant to the solar thermal rocket engine via one or more propellant channels;
introducing a liquid propellant to the solar thermal rocket engine via the one or more propellant channels; and
gradually reducing a gas-to-liquid propellant ratio provided to the solar thermal rocket engine via the one or more propellant channels to ensure that the solar thermal rocket engine is maintained in an instantaneous steady state.
4\. (New) The method of Claim 3, wherein the gas-to-liquid ration is reduced until the gaseous propellant is no longer provided to the solar thermal rocket engine.
5\. (New) A solar thermal rocket engine, comprising:
a combustion chamber;
one or more propellant channels configured to provide propellant to the combustion chamber; and
one or more electric heating elements configured to adjust a temperature of the combustion chamber to an operating temperature.
6\. (New) The solar thermal rocket engine of Claim 5, wherein the one or more electric heating elements are further configured to heat the com combustion chamber to the operating temperature before the one or more propellant channels provide the propellant to the combustion chamber.
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OFFICE ACTION:
DETAILED ACTION
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 . Claims 2-6 are pending in this application.
Claim Objections
Claim 6 is objected to because of the following informalities:
Regarding Claim 6, in Line 2 “com combustion chamber” should be “combustion chamber”.  Appropriate correction is required.
Claim Rejections - 35 USC § 102
The following is a quotation of the appropriate paragraphs of 35 U.S.C. 102 that form the basis for the rejections under this section made in this Office action:
A person shall be entitled to a patent unless –
(a)(1) the claimed invention was patented, described in a printed publication, or in public use, on sale, or otherwise available to the public before the effective filing date of the claimed invention.
(a)(2) the claimed invention was described in a patent issued under section 151, or in an application for patent published or deemed published under section 122(b), in which the patent or application, as the case may be, names another inventor and was effectively filed before the effective filing date of the claimed invention.
Claim(s) 2 is rejected under 35 U.S.C. 102(a)(1) as being anticipated by Kudija (U.S. Patent No. 6,594,984), hereinafter Kudija.
Regarding Independent Claim 2, Kudija discloses solar thermal rocket engine (Figure 1), comprising:
a solar thermal thruster body (Figure 2);
one or more mirrored collectors, 20, configured to concentrate solar energy into an energy beam (Figure 1 – the collector concentrates the solar energy, 101, into a beam); and
an optical aperture, 24 and 25, configured to modulate an amount of the energy beam allowed to pass into the solar thermal thruster body (Figure 4 – the optical aperture is configured such that is modulates the solar flux/ amount of energy that passes into the thruster body based on its orientation).
Claim(s) 5-6 are rejected under 35 U.S.C. 102(a)(1) as being anticipated by Anflo (U.S. Pre-grant Publication 2018/0298846), hereinafter Anflo.
Regarding independent Claim 5, Anflo discloses a solar thermal rocket engine (Figure 3), comprising:
a combustion chamber, 200;
one or more propellant channels, 140, configured to provide propellant to the combustion chamber (the channel provides the propellant to the combustion chamber); and
one or more electric heating elements, 170, configured to adjust a temperature of the combustion chamber to an operating temperature (Paragraph 0049 – the heating element heats the combustion chamber to an operating temperature before firing).
Regarding Claim 6, Anflo discloses the invention as claimed and discussed above. Anflo further discloses the one or more electric heating elements are further configured to heat the combustion chamber to the operating temperature before the one or more propellant channels provide the propellant to the combustion chamber (Paragraph 0049 – the heating element heats the combustion chamber to an operating temperature before firing and thus before the propellant is provided to the combustion chamber).
Allowable Subject Matter
Claims 3-4 are allowed.
Reasons for Allowance
The following is an examiner’s statement of reasons for allowance:
The limitations “gradually reducing a gas-to-liquid propellant ratio provided to the solar thermal rocket engine via the one or more propellant channels to ensure that the solar thermal rocket engine is maintained in an instantaneous steady state” along with the remaining limitations of Claim 3 are not taught or fairly suggested in the prior art of record.
Kokorich (U.S. Pre-grant Publication 2020/00283174) shows a solar thermal rocket that utilizes propellant in different phases (Paragraphs 0004) but is silent regarding gradually reducing a gas-to-liquid propellant ratio provided to the solar thermal rocket engine via the one or more propellant channels to ensure that the solar thermal rocket engine is maintained in an instantaneous steady state.
Claims 4 depend from Claim 3 and necessarily include the allowable subject matter discussed above and therefore are allowed.
Any comments considered necessary by applicant must be submitted no later than the payment of the issue fee and, to avoid processing delays, should preferably accompany the issue fee.  Such submissions should be clearly labeled “Comments on Statement of Reasons for Allowance.”
Conclusion


NEXT OFFICE ACTION: DETAILED ACTION
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 .
Response to Amendment
The amendment filed on 9 June 2025 has been entered. Claim(s) 3-4 and 7-20 remain pending in this application. Claim(s) 1-2 and 5-6 have been cancelled.  Claim(s) 7-20 are new.
Allowable Subject Matter
Claims 3-4 and 7-20 are allowed.
Reasons for Allowance
The following is an examiner’s statement of reasons for allowance:
The limitations “gradually reducing a gas-to-liquid propellant ratio provided to the solar thermal rocket engine via the one or more propellant channels to ensure that the solar thermal rocket engine is maintained in an instantaneous steady state” along with the remaining limitations of Claim 3 are not taught or fairly suggested in the prior art of record.
Kokorich (U.S. Pre-grant Publication 2020/00283174) shows a solar thermal rocket that utilizes propellant in different phases (Paragraphs 0004) but is silent regarding gradually reducing a gas-to-liquid propellant ratio provided to the solar thermal rocket engine via the one or more propellant channels to ensure that the solar thermal rocket engine is maintained in an instantaneous steady state.
Claims 4 and 7-12 depend from Claim 3 and necessarily include the allowable subject matter discussed above and therefore are allowed.
The limitations “gradually reducing a gas-to-liquid propellant ratio provided to the solar thermal rocket engine via the one or more propellant channels to ensure that the solar thermal rocket engine is maintained in an instantaneous steady state” along with the remaining limitations of Claim 7 are not taught or fairly suggested in the prior art of record for substantially the same reasons as discussed above.
Claims 14-20 depend from Claim 13 and necessarily include the allowable subject matter discussed above and therefore are allowed.
Any comments considered necessary by applicant must be submitted no later than the payment of the issue fee and, to avoid processing delays, should preferably accompany the issue fee.  Such submissions should be clearly labeled “Comments on Statement of Reasons for Allowance.”
Conclusion

CITED REFERENCE SUMMARIES: Ref US20180298846A1: Claims:
A reactor Thermal Management System (TMS) for a liquid propellant chemical rocket engine, which system comprises: a reactor for vaporizing a liquid monopropellant and initiating de-composition thereof, said reactor comprising a thermally conductive reactor housing, a catalyst bed, and a heat bed, which is in indirect thermal contact with the catalyst bed via the reactor housing; an electrical heater; and an injector, wherein the heat bed is located upstream of the catalyst bed and downstream of the injector and wherein the injector exhibits one or more openings for the propellant directed laterally towards a portion of the circumference of the inner surface of the reactor housing for injecting the propellant in the form of a cone against said portion, and in the heater being attached to the reactor housing and in thermal contact with said portion, which portion is located upstream of the catalyst bed.
a reactor for vaporizing a liquid monopropellant and initiating de-composition thereof, said reactor comprising a thermally conductive reactor housing, a catalyst bed, and a heat bed, which is in indirect thermal contact with the catalyst bed via the reactor housing;
an electrical heater; and
an injector, wherein the heat bed is located upstream of the catalyst bed and downstream of the injector and wherein the injector exhibits one or more openings for the propellant directed laterally towards a portion of the circumference of the inner surface of the reactor housing for injecting the propellant in the form of a cone against said portion, and in the heater being attached to the reactor housing and in thermal contact with said portion, which portion is located upstream of the catalyst bed.
The reactor TMS of claim 1 , wherein the cone half-angle of the cone is within the range of 35-55Â°.
The reactor TMS of claim 1 , additionally comprising a heat reservoir adjacent to the heater and being in thermal contact with the reactor housing and heater.
The reactor TMS of claim 3 , wherein the heat reservoir is made integral with the reactor housing.
The reactor TMS of claim 3 , wherein the heat reservoir is attached to the outside of the reactor housing.
The reactor TMS of claim 3 , wherein the heat reservoir is attached to the inside of the reactor housing.
The reactor TMS of claim 1 , wherein the injector also exhibits openings for the propellant directed towards the heat bed.
The reactor TMS of claim 1 , comprising a flow restrictor on a propellant feed pipe upstream of the injector and upstream of a propellant flow control valve.
A Liquid Propellant Chemical Rocket Engine (LPCRE), comprising the reactor Thermal Management System of claim 1 , a combustion chamber, and a nozzle.
A spacecraft, satellite, or spaceplane including an engine of claim 9 .
A process of decomposing a liquid propellant in a liquid propellant chemical rocket engine comprising a reactor housing comprising: A subjecting the propellant to a temperature efficient for essentially bringing the propellant into the vapour phase by injecting the propellant towards a heated portion of the inner surface of the reactor housing of the engine, which heated portion is in thermal contact with an electrical heater attached to the reactor housing, which heater is configured to heat said portion; B bringing the essentially vaporised propellant into contact with a catalyst and decomposing the propellant into hot, gaseous combustible components; C combusting the combustible components; and D transferring heat generated by decomposing the propellant and combusting the combustible components to the propellant.
A subjecting the propellant to a temperature efficient for essentially bringing the propellant into the vapour phase by injecting the propellant towards a heated portion of the inner surface of the reactor housing of the engine, which heated portion is in thermal contact with an electrical heater attached to the reactor housing, which heater is configured to heat said portion;
B bringing the essentially vaporised propellant into contact with a catalyst and decomposing the propellant into hot, gaseous combustible components;
C combusting the combustible components; and
D transferring heat generated by decomposing the propellant and combusting the combustible components to the propellant.
The process of claim 11 , wherein, in step D, heat generated in step B and/or in step C is transferred to a heat reservoir, which heat reservoir is in thermal contact with the heater and with said portion of the inner reactor surface towards which portion the propellant is being injected in step A.
The process of claim 11 , wherein the rocket engine comprises a reactor Thermal Management System (TMS), the TMS comprising: a reactor for vaporizing a liquid monopropellant and initiating de-composition thereof, said reactor comprising the thermally conductive reactor housing, a catalyst bed, and a heat bed, which is in indirect thermal contact with the catalyst bed via the reactor housing; the electrical heater; and, an injector, wherein the heat bed is located upstream of the catalyst bed and downstream of the injector and wherein the injector exhibits one or more openings for the propellant directed laterally towards a portion of the circumference of the inner surface of the reactor housing for injecting the propellant in the form of a cone against said portion, and in the heater being attached to and in thermal contact with said portion, which portion is located upstream of the catalyst bed.
a reactor for vaporizing a liquid monopropellant and initiating de-composition thereof, said reactor comprising the thermally conductive reactor housing, a catalyst bed, and a heat bed, which is in indirect thermal contact with the catalyst bed via the reactor housing;
the electrical heater; and,
an injector, wherein the heat bed is located upstream of the catalyst bed and downstream of the injector and wherein the injector exhibits one or more openings for the propellant directed laterally towards a portion of the circumference of the inner surface of the reactor housing for injecting the propellant in the form of a cone against said portion, and in the heater being attached to and in thermal contact with said portion, which portion is located upstream of the catalyst bed.
The process of claim 12 , wherein the heater is configured to heat said portion by direct transfer of heat from said heater.
The process of claim 12 , wherein the heater is configured to heat said portion by transfer of heat from a heat reservoir heated by said heater
The reactor TMS of claim 8 , wherein the flow restrictor comprises a cavitating venturi.
Ref US20200283174A1: Claims:
A method for managing propellant in a spacecraft, the method comprising: storing propellant in a tank as a mixture of liquid and gas; transferring the propellant out of the tank; converting the mixture of liquid and gas propellant into a single phase, where the single phase is either liquid or gaseous; and supplying the single phase of the propellant to a thruster.
storing propellant in a tank as a mixture of liquid and gas;
transferring the propellant out of the tank;
converting the mixture of liquid and gas propellant into a single phase, where the single phase is either liquid or gaseous; and
supplying the single phase of the propellant to a thruster.
The method of claim 1 , wherein converting the propellant into a single phase includes converting the mixture of liquid and gas propellant directly into liquid.
converting the propellant into a single phase includes converting the mixture of liquid and gas propellant directly into liquid.
The method of claim 2 , wherein converting the mixture of liquid and gas propellant directly into liquid includes compressing the propellant using a piston.
converting the mixture of liquid and gas propellant directly into liquid includes compressing the propellant using a piston.
The method of claim 1 , wherein converting the mixture of liquid and gas propellant into the single phase includes converting the propellant directly into gas.
converting the mixture of liquid and gas propellant into the single phase includes converting the propellant directly into gas.
The method of claim 1 , wherein converting the mixture of liquid and gas propellant into a single phase includes: first converting the mixture of liquid and gas propellant into gas, then converting the gas into liquid.
converting the mixture of liquid and gas propellant into a single phase includes: first converting the mixture of liquid and gas propellant into gas, then converting the gas into liquid.
first converting the mixture of liquid and gas propellant into gas, then converting the gas into liquid.
The method of claim 1 , wherein converting the mixture of liquid and gas propellant into the single phase includes drawing the liquid through a wicking material into a liquid-phase pump.
The method of claim 6 , further comprising: causing a low-density vapor-phase of the propellant to condense onto a cooled complex surface at a temperature below a dew point of the propellant but above the freezing point of the propellant.
causing a low-density vapor-phase of the propellant to condense onto a cooled complex surface at a temperature below a dew point of the propellant but above the freezing point of the propellant.
The method of claim 6 , wherein the complex surface is composed of loosely packed hydrophilic fibers.
The method of claim 6 , further comprising: progressively compressing, using a peristaltic-type pump, the wicking material along a wave which forces liquid out of the wicking material and toward an output port.
progressively compressing, using a peristaltic-type pump, the wicking material along a wave which forces liquid out of the wicking material and toward an output port.
The method of claim 1 , wherein converting the mixture of liquid and gas propellant into the single phase includes: evaporating the mixture to a complete vapor phase; and condensing the complete vapor phase to a complete liquid phase.
evaporating the mixture to a complete vapor phase; and
condensing the complete vapor phase to a complete liquid phase.
The method of claim 10 , wherein the evaporating includes: directing a stream of the mixture through a restrictive orifice to generate an abrupt pressure drop to flash-evaporate the mixture.
directing a stream of the mixture through a restrictive orifice to generate an abrupt pressure drop to flash-evaporate the mixture.
The method of claim 11 , wherein the condensing includes: adding heat to the mixture using a warmed evaporator to generate a low-pressure vapor; compressing a stream of the low-pressure vapor to generate a pressurized vapor with a higher pressure using a vapor pump; and condensing the pressurized vapor to the complete liquid phase using a cooled condenser.
adding heat to the mixture using a warmed evaporator to generate a low-pressure vapor;
compressing a stream of the low-pressure vapor to generate a pressurized vapor with a higher pressure using a vapor pump; and
condensing the pressurized vapor to the complete liquid phase using a cooled condenser.
The method of claim 12 , further comprising: transferring heat from the cooled condenser to the warmed evaporator.
transferring heat from the cooled condenser to the warmed evaporator.
The method of claim 10 , wherein evaporating the mixture to the complete vapor phase includes passing the mixture in pulses through a fast-acting valve into a low-pressure chamber.
A spacecraft comprising: a tank storing propellant in a tank as a mixture of liquid and gas; a thruster configured to consume the propellant to generate thrust; a two-phase intake component configured to configured to receive the mixture of liquid and gas from the tank; and a phase conversion component configured to (i) receive the mixture of liquid and gas from the two-phase intake component, (ii) convert the mixture of liquid and gas into a single phase of the propellant, and (iii) supply the single phase of the propellant to the thruster.
a tank storing propellant in a tank as a mixture of liquid and gas;
a thruster configured to consume the propellant to generate thrust;
a two-phase intake component configured to configured to receive the mixture of liquid and gas from the tank; and
a phase conversion component configured to (i) receive the mixture of liquid and gas from the two-phase intake component, (ii) convert the mixture of liquid and gas into a single phase of the propellant, and (iii) supply the single phase of the propellant to the thruster.
The spacecraft of claim 15 , wherein: the two-phase intake component includes porous wicking material to absorb the mixture; and the phase conversion component includes a mechanism configured to compress the porous wicking material to thereby extract a liquid phase of the propellant.
the two-phase intake component includes porous wicking material to absorb the mixture; and
the phase conversion component includes a mechanism configured to compress the porous wicking material to thereby extract a liquid phase of the propellant.
The method of claim 16 , wherein: the phase conversion component further includes a cooled complex surface, and the phase conversion component causes a low-density vapor-phase of the propellant to condense onto the cooled complex surface at a temperature below a dew point of the propellant but above the freezing point of the propellant.
the phase conversion component further includes a cooled complex surface, and
the phase conversion component causes a low-density vapor-phase of the propellant to condense onto the cooled complex surface at a temperature below a dew point of the propellant but above the freezing point of the propellant.
The spacecraft of claim 17 , wherein the complex surface is composed of loosely packed hydrophilic fibers.
The spacecraft of claim 15 , wherein: the phase conversion component includes a piston configured to convert the mixture of liquid and gas propellant directly into liquid.
the phase conversion component includes a piston configured to convert the mixture of liquid and gas propellant directly into liquid.
The spacecraft of claim 15 , wherein the phase conversion component is configured to: evaporate the mixture to a complete vapor phase; and condense the complete vapor phase to a complete liquid phase.
evaporate the mixture to a complete vapor phase; and
condense the complete vapor phase to a complete liquid phase.
Ref US6594984B1: [content unavailable]
Ref US6594984: Claims:
1. A combined generator device for use in a satellite, said combined generator device capable of generating electrical power and thrust from an optical solar image including a high-intensity region and a low-intensity region, wherein the high-intensity region has a higher radiant solar energy intensity than the low-intensity region, said combined generator device comprising:
a solar concentrator configured to focus and direct the regions of the optical image;
a first receiver including a first radiant energy absorbing member positioned so as to receive the focused low-intensity region of the optical solar image, the receiver further comprising a thermal storage medium thermally coupled to the first radiant energy absorbing member to store thermal energy absorbed thereby;
an electrical power generator thermally coupled to the first receiver and receiving thermal energy therefrom, the electrical power generator configured to convert the thermal energy into electrical power; and
a thrust generator having a second receiver and a propellant supply, the second receiver including a second radiant energy absorbing member positioned so as to receive the focused high-intensity region of the optical solar image, the propellant supply thermally coupled with the second radiant energy absorbing member and receiving thermal energy therefrom so as to generate thermal propulsion.
2. A combined generator device of claim 1, wherein the solar concentrator includes a concave mirror and a pointing system supporting the mirror, wherein the pointing system is capable of repositioning the mirror so as to reflect and focus the optical solar image onto the receivers.
3. A combined generator device of claim 1, wherein the first radiant energy absorbing member defines a first cavity having an aperture positioned to receive the low-intensity region and wherein the second radiant energy absorbing member of the second receiver is positioned within the first cavity to receive the region of high-intensity.
4. A combined generator device of claim 3, wherein the first cavity defined by the first radiant energy absorbing member is a cylindrical cavity and wherein the second receiver is concentrically positioned within the first cavity.
5. A combined generator device of claim 4, wherein the thermal energy storage medium is layered around an outer surface of the first radiant energy absorbing member.
6. A combined generator device of claim 5, wherein the thermal energy storage medium is a solid medium.
7. A combined generator device of claim 5, wherein the thermal energy storage medium is constructed of a graphite material.
8. A combined generator device of claim 4, wherein the second radiant energy absorbing member is constructed of a refractory material.
9. A combined generator device of claim 8, wherein the refractory material is rhenium.
10. A combined generator device of claim 4, wherein the second radiant energy absorbing member defines a second cavity having an aperture.
11. A combined generator device of claim 10, wherein the thrust generator further includes a heat exchanger thermally coupled with the second radiant energy absorbing member, said heat exchanger connected in fluid communication with the propellant supply and heating propellant passing therethrough.
12. A combined generator device of claim 11, wherein the thrust generator further includes a pre-heat exchanger thermally coupled with the first radiant energy absorbing member, said pre-heat exchanger connected in fluid communication with the propellant supply upstream of the heat exchanger and pre-heating propellant passing therethrough.
13. A combined generator device of claim 12, wherein the thrust generator further includes a thrust nozzle in fluid communication with the propellant supply downstream from the heat exchanger.
14. A combined generator device of claim 1, wherein the solar concentrator is configured to focus the optical solar image onto a focal plane and wherein the first and second energy absorbing members are positioned at the focal plane.
15. A solar power receiver assembly for converting an optical solar image into thermal energy for use by a thrust generator and an electrical power generator, said optical solar image including a high-intensity region and a low-intensity region wherein the high-intensity region has a higher radiant solar energy intensity than the low-intensity region, said combined generator device comprising:
a solar concentrator configured to focus and direct the regions of the optical image;
a first receiver including a first radiant energy absorbing member positioned so as to receive the focused low-intensity region of the optical solar image, the receiver further comprising a thermal storage medium thermally coupled to the first radiant energy absorbing member to store thermal energy absorbed thereby; and
a second receiver including a second radiant energy absorbing member positioned so as to receive the focused high-intensity region of the optical solar image wherein said electrical power generator is thermally coupled to the first receiver and converts thermal energy from the first receiver into electrical power and wherein said thrust generator is thermally coupled to the second receiver and uses thermal energy from the second receiver to produce thrust.
16. A combined generator device of claim 15, wherein the solar concen
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