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 .
Claim Rejections - 35 USC § 112
The following is a quotation of the first paragraph of 35 U.S.C. 112(a):
(a) IN GENERAL.—The specification shall contain a written description of the invention, and of the manner and process of making and using it, in such full, clear, concise, and exact terms as to enable any person skilled in the art to which it pertains, or with which it is most nearly connected, to make and use the same, and shall set forth the best mode contemplated by the inventor or joint inventor of carrying out the invention.
Claims 21-40 are rejected under 35 U.S.C. 112(a) or 35 U.S.C. 112 (pre-AIA ), first paragraph, as failing to comply with the written description requirement. The claim(s) contains subject matter which was not described in the specification in such a way as to reasonably convey to one skilled in the relevant art that the inventor or a joint inventor, or for applications subject to pre-AIA 35 U.S.C. 112, the inventor(s), at the time the application was filed, had possession of the claimed invention.
Claims 21, 33 recite conflicting requirements from different embodiments and applicant did not have possession of the limitations of claims 21, 33 in the same embodiment. Applicant’s claims have been amended to not read on any single embodiment and applicant did not have possession of these limitations together – see paragraph b of the 112(b) rejection below, which will not be repeated for conciseness.
The following is a quotation of 35 U.S.C. 112(b):
(b) CONCLUSION.—The specification shall conclude with one or more claims particularly pointing out and distinctly claiming the subject matter which the inventor or a joint inventor regards as the invention.
Claims 21-40 are rejected under 35 U.S.C. 112(b) or 35 U.S.C. 112 (pre-AIA ), second paragraph, as being indefinite for failing to particularly point out and distinctly claim the subject matter which the inventor or a joint inventor (or for applications subject to pre-AIA 35 U.S.C. 112, the applicant), regards as the invention.
Claims 21, 33 recite respectively: (21) “wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow; and an exhaust section, a gas flowpath defined and extending through the inlet section, the combustion section, and the exhaust section, the gas flowpath extending at least partially along a centerline of the combustion section, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, wherein the combustion section includes:(a) a first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with the first portion of the airflow from the inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor”.
(33) “wherein an airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, the gas flowpath extending at least partially along a centerline of the combustion section, the first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with the first portion of the airflow received by the first combustion portion from an inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor”
It is unclear if “the first portion of the airflow from the inlet section” in paragraph (a) of claim 22 and in analogous portion of claim 33 requires both “a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow” to be delivered to the first combustion portion or could be just one of them.
Second, it is unclear which reference point applicant uses for to determine where “radially inner” “radially outer” are determined for “wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow.” Applicant argues relative to Fig. 4, but that Figure when applied to most of the disclosed Figs of the actual engine, e.g. Fig. 1a, 1b, 2b, 7A, etc only have a single first flow 81(a). In other words, there is no “a radially inner stream of the first portion of the airflow” present. In other words, Fig. 4 specifically only shows a first single flow 81(a) and does not show a radially inner and radially outer first flow with the second flow in between. Even for embodiments [e.g. Fig. 2A] that have a radially inner stream of the first portion of the airflow [to 125(b)] vs the first radially outer stream of the first portion [to 125(a)] of the airflow, these first outer and inner portions of air do not flow to the same first combustion portions -- 125(b) and (125a) are different combustion regions. Accordingly, applicant’s claims are unclear because applicant appears to mix and match the embodiments in an incompatible fashion. If applying a first engine construction [e.g. Fig. 4 with any of Figs. 1a, 1b, 2b, 7A], then there is no radially inner stream of the first portion of the airflow – but only a single first portion of airflow. When applying the second type of engine construction [e.g. Fig. 2A], the first portion of the airflows do not flow to the same first combustion portion, as 125(a) and 125(b) are different combustion portions. Applicant’s claims do not read on his own invention and are thus 1) indefinite and 2) new matter as applicant did not possess this juxtaposition of claim limitations addressed above.
Claims 21 & 33 are rejected under 35 U.S.C. 112(b) or 35 U.S.C. 112 (pre-AIA ), second paragraph, as being incomplete for omitting essential structural cooperative relationships of elements, such omission amounting to a gap between the necessary structural connections. See MPEP § 2172.01. The omitted structural cooperative relationships are: there is no way to ascertain where “wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow” is located. There is nothing in the claim to structurally separate the portions of the airflow and determine the radially contiguous location(s). Moreover, this limitation is unclear because applicant has attempted to define separate airflows, but air flows where it will and “a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow” is unclear because even some of the first airflow 81(a) may split and go both the first combustion portion and the second combustion portion [see annotations]. Similarly, some of the second airflow may split and go to both the first combustion portion and the second combustion portion.
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Claim Interpretation
Claims 21, 24-30 are generally parallel to claims 33-40. For conciseness, if claim limitations were already treated in one set of claims, they may not be repeated for the other set.
Claim Rejections - 35 USC § 103
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.
Claim(s) 20-40, as understood, is/are rejected under 35 U.S.C. 103 as being unpatentable over CN 102588145 in view of CN 104948348.
“In the above-mentioned accompanying drawing: 1 air, 2 premixed gas, 3 detonation combustion gas, 4 rotation detonation wave ejectors, 5 lobe nozzles with injectors, 6 interior ram air duct; 7 outer shroud lobe nozzles with injectors, the multi-modal detonation combustion of 8 expandings chamber, 9 expanding jet pipes, 10 vortex intake ducts, 11 outer shroud casings, 12 outer shroud ram air duct; 13 rotation detonation gas outlets, 14 annular phase place detonation wave control system, 15 annular phase place detonation waves, 16 jet pipe venturis, 17 phase vectors detonation waves; 18 fuel oil inlet pipes, 19 fuel oil outlet tubes, 20 ultrasonic generators, 21 semicircle pipeline fluidized beds, 22 fuel pumps; 23 fuel oil ejectors, 24 fuel oil fuel tanks, 25 catalyzer, 26 injection intakepories, 27 intake ducts. [see page 3 of the translation]”
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CN ‘145 teaches A supersonic engine comprising; an inlet section 1 to which an inlet airflow is provided; a combustion section that receives an airflow from the inlet section, wherein the airflow received by the combustion section includes a first portion [see annotations] and a second portion [see annotations – but central flow], and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow; and an exhaust section 14, a gas flowpath defined extending through the inlet section, the combustion section, and the exhaust section, the gas flowpath extending at least partially along a centerline of the combustion section [along leadlines for 1 to 3], wherein both the first portion and the second portion of the airflow are located within the open flowpath, wherein the combustion section includes (a) a first combustion portion 4 having a rotating detonation combustor to which a first flow of fuel is provided to mix with airflow to generate a first fuel-air mixture 4 in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor, and (b) a second combustion portion [downstream of 5 or 7] arranged downstream of the rotating detonation combustor 4, a second flow of fuel [central 2, where 2 is fuel/air mixture] being provided to the second combustion portion to mix with the second portion of the airflow 2 from the inlet section and with combustion products from the rotating detonation combustor 4 to generate a second fuel-air mixture 2, the combustion products from the rotating detonation combustor 4 functioning as a pilot to ignite the second fuel-air mixture 2 [central 2, note there is no separate igniter for the downstream combustion and the top of page 3 of the machine translation teaches the initiation / ignition of the premixed gas 2 by the injection shear layer; in other words the hot combustion gases from 4 serves as a pilot / ignites the downstream fuel / air mixtures 2, 3];
(33) A combustion section for a supersonic engine, comprising: a first combustion portion and a second combustion portion defining a gas flowpath therethrough, wherein the airflow received by the combustion section includes a first portion [see annotations] and a second portion [see annotations – but central flow], and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, the gas flowpath extending at least partially along a centerline of the combustion section [along leadlines for 1 to 3], the first combustion portion having a rotating detonation combustor 4 to which a first flow of fuel is provided to mix with airflow received by the first combustion portion to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor; and the second combustion portion [downstream of 5 or 7] arranged downstream of the rotating detonation combustor, a second flow of fuel [2 is fuel/air mixture] being provided to the second combustion portion to mix with the second portion of the airflow from the inlet section and combustion products from the rotating detonation combustor 4 to generate a second fuel-air mixture [downstream of 5 or 7], the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture; (22) wherein the supersonic engine is a ramjet engine and the second combustion portion is one of a ramjet combustor or an augmenter [ram air is taught and at subsonic conditions, scramjets are operable as ramjets. Alternately, it would have been obvious to operate as a ramjet engine]; (23) wherein the supersonic engine is a scramjet engine and the second combustion portion is one of a scramjet combustor or an augmenter [scramjet – middle of page 1]; (24, 34) wherein the combustion section includes a longitudinal wall extending along a length of the combustion section and defining an outer periphery of the first combustion portion and the second combustion portion, the gas flowpath being defined within the outer periphery of the combustion portion and the second combustion portion; wherein the rotating detonation combustor extends about the outer periphery of the first combustion portion; (26, 36) wherein the first combustion portion includes an inner wall assembly [see annotations] extending from the longitudinal wall into the gas flowpath, the inner wall assembly having an upstream face [upstream edge of lobe mixer 5] and a downstream wall;(27, 37) wherein the inner wall assembly, downstream of the first combustion portion, extends about a periphery of the longitudinal wall and defines the gas flowpath therethrough
wherein a rotating detonation region 4 is defined at the upstream face [lobe mixer 5] of the inner wall assembly; wherein the upstream face of the inner wall assembly captures the first portion of the airflow from the inlet section; (30, 40) wherein the upstream face of the inner wall assembly includes an inner tip 5 extending into the gas flowpath, and the downstream wall extends from the inner tip to the longitudinal wall; wherein the combustion products from the rotating detonation combustor flow over the inner tip into the gas flowpath to mix with the second portion of the airflow; wherein the downstream wall includes at least one second fuel injection port [e.g. from 7], and the second flow of fuel 2 is provided into the second combustion portion via the at least one second fuel injection port 7. CN ‘145 does not teach the air 10 and combusting in 4 is from the first portion of the airflow from the inlet section [for one of the interpretations, for the other interpretation, the first portion of airflow in 12 is from the inlet]. CN ‘348 [Figs. 2, 3] teaches a combustion section that receives an airflow from the inlet section 3, wherein the combustion section defines a central axis, wherein the airflow includes a first portion 3, 16, 9 and a second portion 3, 6, and wherein the second portion of the airflow 3, 6 is radially inward of the first portion 3, 16, 9 of the airflow and flowing along the central axis; and the first portion of the airflow 3, 16, 9 from the inlet section is delivered to the combustor and the second portion of the airflow 3, 6 is delivered to the combustor. Alternately, CN ‘348 [Fig. 3] teaches wherein the combustion section defines a central axis, wherein the airflow includes a first portion 3, 13, 14, 15 and a second portion 3, 6 and wherein the second portion of the airflow 3, 6 is radially inward of the first portion 3, 13, 14, 15 of the airflow and flowing along the central axis; the first portion of the airflow 3, 13, 14, 15 from the inlet section is delivered to the combustor and the second portion of the airflow 3, 6 is delivered to the combustor. Yet another alternative is that the first [inner and outer] and second portions of airflow may be construed as annotated. It would have been obvious to one of ordinary skill in the art to provide the first portion of the airflow from the inlet section the to mix with the first fuel of CN ’ 145, in the manner taught by CN ‘348, in order to facilitate multiple flows of air into the combustor and/or provide a convenient source of oxygen / air for the combustor, as typically done in the art.
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Claim(s) 21-40, as understood, is/are rejected under 35 U.S.C. 103 as being unpatentable over Morrison et al (3,240,010) in view of any of Townend (3,603,094), Rich (2,952,123), Pavia et al (2011/0005193) and further in view of Andreadis (2008/0196414).
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Morrison teaches A supersonic engine [intended use, note rockets and ramjets are capable of supersonic operation] comprising; an inlet section [16 or upstream of 16 when applied to air breathing ramjet; col. 6, lines 62-72] to which an inlet airflow is provided; a combustion section 17 that receives an airflow from the inlet section; wherein the airflow received by the combustion section includes a first portion and a second portion [see annotations], and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, and an exhaust section 50, a gas flowpath defined extending through the inlet section, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is contiguous [adjacent at any point in space] with and radially inward of the first portion of the airflow [ramjet air flows into the combustor in an air breathing ramjet; col. 6, lines 62-72], the combustion section, and the exhaust section, the gas flowpath extending at least partially along a centerline of the combustion section therewithin, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, wherein the combustion section includes (a) a first combustion portion having a rotating detonation combustor 17 to which a first flow of fuel is provided to mix with a first portion of the airflow from the inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor, and
(33) A combustion section for a supersonic engine [intended use, note rockets and ramjets are capable of supersonic operation], comprising: a first combustion portion 17 defining a gas flowpath therethrough, wherein the airflow received by the combustion section includes a first portion and a second portion [see annotations], and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, the first combustion portion having a rotating detonation combustor 17 to which a first flow of fuel is provided to mix with a first portion of an airflow received by the first combustion portion from an inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor; (24, 34) wherein the combustion section includes a longitudinal wall 20 extending along a length of the combustion section and defining an outer periphery of the first combustion portion and the second combustion portion, the gas flowpath being defined within the outer periphery of the combustion portion and [exhaust portion]; wherein the rotating detonation combustor 17 extends about the outer periphery of the first combustion portion; (26, 36) wherein the first combustion portion includes an inner wall assembly 18 extending from the longitudinal wall into the gas flowpath, the inner wall assembly having an upstream face 18 and a downstream wall 40; wherein the inner wall assembly, downstream of the first combustion portion, extends about a periphery of the longitudinal wall and defines the gas flowpath therethrough;
wherein a rotating detonation region is defined at the upstream face 18 of the inner wall assembly; wherein the upstream face of the inner wall assembly captures the first portion of the airflow from the inlet section; (30, 40) wherein the upstream face of the inner wall assembly includes an inner tip 48 extending into the gas flowpath, and the downstream wall extends from the inner tip to the longitudinal wall; wherein the combustion products from the rotating detonation combustor flow over the inner tip into the gas flowpath to mix with the second portion of the airflow [inner airflow].
First beyond the intended use, it is noted regarding the “supersonic engine”, it is well known in the art to use ramjets / rockets as supersonic engines and it would have been obvious to do so as typical range of operation utilized in the art – alternately this is taught by Andreadis in the scramjet treatment. Morrison does not necessarily illustrated the second portion of the airflow is radially inward of the first portion of the airflow and flowing along the central axis [though this is applied as illustrated, noting that this limitation only requires a direction of flow axially and has not been construed as flowing on the axis, as that would already disqualify many of the applicant’s actual illustrated engine embodiments as generally addressed in paragraph 9 of the final rejection of 4/22/2025]. Andreadis teach wherein the combustion section [Figs. 5, 6] defines a central axis, wherein the airflow includes a first portion [outer] and a second portion [central axis], and wherein the second portion of the airflow is radially inward of the first portion of the airflow and flowing along the central axis and is consistent with the desire of Morrison to use his rotating detonation chamber with ramjets, where the inlet is not illustrated but disclosed. It would have been obvious to one of ordinary skill in the art to make the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and flowing along the central axis, in the manner taught by Andreadis, as the typical flow of air within the ramjet engine.
Morrison does not teach a second combustion portion, a second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with the second portion of the airflow from the inlet section and combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture; nor the downstream wall includes at least one second fuel injection port, and the second flow of fuel is provided into the second combustion portion via the at least one second fuel injection port. Townend teaches a second combustion portion 12, a second combustion portion 12 arranged downstream of the first combustor [upstream of 12], a second flow of fuel 15 being provided to the second combustion portion to mix with the second portion of the airflow from the inlet section and combustion products from the first combustor to generate a second fuel-air mixture, the combustion products from the first combustor functioning as a pilot to ignite the second fuel-air mixture 15 [fuel 15 is ignited by the hot exhaust gas at B, see paragraph bridging cols. 2 and 3]; and the downstream wall includes at least one second fuel injection port 15, and the second flow of fuel is provided into the second combustion portion via the at least one second fuel injection port 15. The second fuel and second oxidizer [see col. 2, lines 1-25] are used to reduce noise and increase thrust as well reduce drag at the exit of the nozzle [see also col. 1, lines 53-63]. Rich teaches a second combustion portion, a second combustion portion 23, 17 [Fig. 6] arranged downstream of the first combustor [from 10], a second flow of fuel 23, 17 [Fig. 6]being provided to the second combustion portion to mix with the second portion of the airflow from the inlet section and combustion products from the first combustor [from 10] to generate a second fuel-air mixture, the combustion products from the first combustor functioning as a pilot to ignite the second fuel-air mixture [inherent, as there is no igniter for the second fuel disclosed]; and the downstream wall includes at least one second fuel injection port 23, 1, and the second flow of fuel is provided into the second combustion portion via the at least one second fuel injection port. Rich teaches burning the second flow of fuel downstream of the combustion gases from 10 increases the thrust of the jet engine and/or is used for thrust vectoring [col. 5, lines 27+]. Pavia et al teach a second combustion portion 166, a second combustion portion 166 arranged downstream of the first combustor 158, 122, a second flow of fuel 128 being provided to the second combustion portion to mix with the second portion of the airflow from the inlet section and combustion products from the first combustor to generate a second fuel-air mixture, the combustion products from the first combustor functioning as a pilot to ignite the second fuel-air mixture [inherent, no igniter taught]; and the downstream wall 162 includes at least one second fuel injection port 128, and the second flow of fuel is provided into the second combustion portion via the at least one second fuel injection port. Pavia et al teach
“[0029] 3.) the turbulence 166 caused increased mixing and combustion between the fuel film coolant and the core combustion gases 158 in the expansion nozzle 180 reducing the performance losses of the combustion device. [paragraph 0029] "
Pavia et al also teach the fuel cools the rocket nozzle walls and therefore prolongs its life. It would have been obvious to one of ordinary skill in the art to employ a second combustor and a second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with the second portion of the airflow from the inlet section and combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture, using the second fuel injection in the diverging section / downstream face 44 of Morrison, as taught by any of Townend, Rich and Pavia et al, in order to reduce noise and increase thrust as well as reduce drag at the exit of the nozzle and/or to increase the thrust of the jet engine or permit used for thrust vectoring and/or to cool the nozzle while reducing performance losses. In combination, the prior art teach (22) wherein the supersonic engine is a ramjet engine [Morrison] and the second combustion portion is one of a ramjet combustor or an augmenter. Furthermore, as for (23) wherein the supersonic engine is a scramjet engine and the second combustion portion is one of a scramjet combustor or an augmenter, Morrison does not teach specifically using a scramjet. Andreadis teaches the use of a supersonic engine with dual mode ramjet and scramjet [with subsonic and supersonic combustion – see paragraph 0002]. It would have been obvious to one of ordinary skill in the art to employ the combustor of Morrison in a scramjet, i.e. with supersonic combustion or ramjet or as part of a supersonic engine, as an Andreadis teaches using the same combustor structure in both supersonic ramjets and scramjets. The second combustion chamber added by any of Townend, Rich and Pavia, will be either an augmenter or ramjet / scramjet combustor. Note that Townend specifically teaches the second combustor is piloted by the first combustor, and Rich and Pavia do not teach an igniter is necessary and thus inherently piloted by the hot combustion gases, in the manner of Townend. Alternately, Andreadis teach the second fuel 108 is piloted by the first combustor 38 [e.g. paragraph 0026]. It would have been obvious to one of ordinary skill in the art to make the first combustor function as a pilot to ignite the second fuel-air mixture, as taught by either Townend or Andreadis, as the typical practice in the art that promotes simplicity of operation by making the second combustor ignitorless.
Double Patenting
The nonstatutory double patenting rejection is based on a judicially created doctrine grounded in public policy (a policy reflected in the statute) so as to prevent the unjustified or improper timewise extension of the “right to exclude” granted by a patent and to prevent possible harassment by multiple assignees. A nonstatutory double patenting rejection is appropriate where the conflicting claims are not identical, but at least one examined application claim is not patentably distinct from the reference claim(s) because the examined application claim is either anticipated by, or would have been obvious over, the reference claim(s). See, e.g., In re Berg, 140 F.3d 1428, 46 USPQ2d 1226 (Fed. Cir. 1998); In re Goodman, 11 F.3d 1046, 29 USPQ2d 2010 (Fed. Cir. 1993); In re Longi, 759 F.2d 887, 225 USPQ 645 (Fed. Cir. 1985); In re Van Ornum, 686 F.2d 937, 214 USPQ 761 (CCPA 1982); In re Vogel, 422 F.2d 438, 164 USPQ 619 (CCPA 1970); In re Thorington, 418 F.2d 528, 163 USPQ 644 (CCPA 1969).
A timely filed terminal disclaimer in compliance with 37 CFR 1.321(c) or 1.321(d) may be used to overcome an actual or provisional rejection based on nonstatutory double patenting provided the reference application or patent either is shown to be commonly owned with the examined application, or claims an invention made as a result of activities undertaken within the scope of a joint research agreement. See MPEP § 717.02 for applications subject to examination under the first inventor to file provisions of the AIA as explained in MPEP § 2159. See MPEP § 2146 et seq. for applications not subject to examination under the first inventor to file provisions of the AIA . A terminal disclaimer must be signed in compliance with 37 CFR 1.321(b).
The filing of a terminal disclaimer by itself is not a complete reply to a nonstatutory double patenting (NSDP) rejection. A complete reply requires that the terminal disclaimer be accompanied by a reply requesting reconsideration of the prior Office action. Even where the NSDP rejection is provisional the reply must be complete. See MPEP § 804, subsection I.B.1. For a reply to a non-final Office action, see 37 CFR 1.111(a). For a reply to final Office action, see 37 CFR 1.113(c). A request for reconsideration while not provided for in 37 CFR 1.113(c) may be filed after final for consideration. See MPEP §§ 706.07(e) and 714.13.
The USPTO Internet website contains terminal disclaimer forms which may be used. Please visit www.uspto.gov/patent/patents-forms. The actual filing date of the application in which the form is filed determines what form (e.g., PTO/SB/25, PTO/SB/26, PTO/AIA /25, or PTO/AIA /26) should be used. A web-based eTerminal Disclaimer may be filled out completely online using web-screens. An eTerminal Disclaimer that meets all requirements is auto-processed and approved immediately upon submission. For more information about eTerminal Disclaimers, refer to www.uspto.gov/patents/apply/applying-online/eterminal-disclaimer..
Claims 21-40 are rejected on the ground of nonstatutory double patenting as being unpatentable over claims 1-17 of U.S. Patent No. 11486579 and especially in view of Johnson et al (2,966,028) and optionally Morrison et al (3,240,010) and Andreadis et al (2008/0196414) and any of the prior art, as applied above [e.g. CN ‘145]. The other prior art applied above are also applicable for the dependent claims and are incorporated from above for the various combinations involving Morrison or CN’145 and not repeated for conciseness.
US 11486579 teaches the majority of the claimed limitations with minor variations in claim wording (e.g. exhaust section vs exhaust nozzle) and include an inlet section in claim 12 but do not clearly teach the first and second airflows are from the inlet section to generate the respective first and second fuel/air mixtures and wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and flowing along the central axis, wherein both the first portion and the second portion of the airflow are located within the gas flowpath. This is well known from the prior art of record applied above, particularly in light of Andreadis, to use the airflow from the inlet to supply the air from the air source to the gas flowpath which becomes the claimed first and second portions of airstream and wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and flowing along the central axis. It would have been obvious to one of ordinary skill in the art to employ an inlet to provide the claimed first and second portions of airstream to the first and second combustion regions, wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and flowing along the central axis, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, as an obvious type of delivery system for the combustible gases needed to form first and second combustion regions. As for the supersonic engine or scramjets, this is taught by CN ‘145 or Andreadis. This would be obvious as a typical use of rotating detonation and/or ramjets / scramjets. As for the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture, this is well known in the prior art Morrison does not teach specifically using a scramjet. As for the limitation, wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, this is taught by e.g. applied prior art and incorporated by reference from the prior art treatment above. It would have been obvious to one of ordinary skill in the art to employ wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, in the manner applied above, as an obvious flow pattern used in the art for airflow. Andreadis teaches the use of dual mode ramjets and scramjets [with subsonic and supersonic combustion – see paragraph 0002]. It would have been obvious to one of ordinary skill in the art to employ the combustor of Morrison in a scramjet, i.e. with supersonic combustion, as an Andreadis teaches using the same combustor structure in both ramjets and scramjets. The second combustion chamber added by any of Townend, Rich and Pavia, will be either an augmenter or ramjet / scramjet combustor. Note that Townend specifically teaches the second combustor is piloted by the first combustor, and Rich and Pavia do not teach an igniter is necessary and thus inherently piloted by the hot combustion gases, in the manner of Townend. Alternately, Andreadis teach the second fuel 108 is piloted by the first combustor 38 [e.g. paragraph 0026]. It would have been obvious to one of ordinary skill in the art to make the first combustor function as a pilot to ignite the second fuel-air mixture, as taught by either Townend or Andreadis, as the typical practice in the art that promotes simplicity of operation by making the second combustor ignitorless. As for the supersonic combusting ramjet, Morrison teaches a ramjet with rotating detonation and Andreadis teaches the supersonic combusting type of ramjet are typical types of ramjets used in the art. It would have been obvious to one of ordinary skill in the art to employ supersonic combustion in the ramjet/scramjet, as taught by Andreadis, in order to be operable at high flight speeds. Additionally, the other claim limitations are taught by the applied prior art, as applied in the various Morrison combinations, and are not repeated for conciseness. For claim 33, a change in preamble from a combustion section vs an engine with claimed combustion section, is an obvious variation from an engine with the claimed combustion section. It would have been obvious to employ the combustion section for a supersonic engine as a typical application of the claimed patent combustor.
18643376
US 11486579
21. A supersonic engine comprising;
an inlet section to which an inlet airflow is provided; a combustion section that receives an airflow from the inlet section, wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and flowing along the central axis; and an exhaust section [exhaust nozzle in ‘579 patent], a gas flowpath defined extending through the inlet section, the combustion section, and the exhaust section, the gas flowpath extending at least partially along a centerline of the combustion section, wherein both the first portion and the second portion of the airflow are located within the gas flowpath,
wherein the combustion section includes (a) a first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with a first portion of the airflow from the inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor, and
(b) a second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with a second portion of the airflow from the inlet section and with combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture.
22. The supersonic engine according to claim 21, wherein the supersonic engine is a ramjet engine and the second combustion portion is one of a ramjet combustor or an augmenter.
23. The supersonic engine according to claim 21, wherein the supersonic engine is a scramjet engine and the second combustion portion is one of a scramjet combustor or an augmenter.
24. The supersonic engine according to claim 21, wherein the combustion section includes a longitudinal wall extending along a length of the combustion section and defining an outer periphery of the first combustion portion and the second combustion portion, the gas flowpath being defined within the outer periphery of the combustion portion and the second combustion portion.
25. The supersonic engine according to claim 24, wherein the rotating detonation combustor extends about the outer periphery of the first combustion portion.
26. The supersonic engine according to claim 25, wherein the first combustion portion includes an inner wall assembly extending from the longitudinal wall into the gas flowpath, the inner wall assembly having an upstream face and a downstream wall.
27. The supersonic engine according to claim 26, wherein the inner wall assembly surrounds the outer periphery of the first combustion portion and defines the gas flowpath therethrough.
28. The supersonic engine according to claim 26, wherein a rotating detonation region is defined at the upstream face of the inner wall assembly.
29. The supersonic engine according to claim 26, wherein the upstream face of the inner wall assembly captures the first portion of the airflow from the inlet section.
30. The supersonic engine according to claim 26, wherein the upstream face of the inner wall assembly includes an inner tip extending into the gas flowpath, and the downstream wall extends from the inner tip to the longitudinal wall.
31. The supersonic engine according to claim 30, wherein the combustion products from the rotating detonation combustor flow over the inner tip into the gas flowpath to mix with the second portion of the airflow.
32. The supersonic engine according to claim 30, wherein the downstream wall includes at least one second fuel injection port, and the second flow of fuel is provided into the second combustion portion via the at least one second fuel injection port.
33. A combustion section for a supersonic engine, comprising: a first combustion portion and a second combustion portion defining a gas flowpath therethrough, wherein the combustion section defines a central axis, wherein an airflow into the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first, the first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with a first portion of an airflow received by the first combustion portion from an inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor; and the second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with a second portion of the airflow from the inlet section and with combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture.
1. An engine, the engine comprising: a longitudinal wall extended along a lengthwise direction, wherein the longitudinal wall defines a gas flowpath of the engine; and an inner wall assembly extended inward from the longitudinal wall into the gas flowpath,
wherein the inner wall assembly defines a rotating detonation combustion region in a first portion of the gas flowpath upstream of the inner wall assembly and adjacent to the longitudinal wall along the lengthwise direction and extending inward a depth from the longitudinal wall, a second portion of the gas flowpath being surrounded by the rotating detonation region and extending in the lengthwise direction from upstream of the rotating detonation region to downstream of the rotating detonation region into an additional combustion region downstream of the rotating detonation region, wherein the inner wall assembly comprises (a) a concave upstream face extended inward from the longitudinal wall into the gas flowpath adjacent to the rotating detonation combustion region, and (b) a downstream face extending at least partially along the lengthwise direction from an inward-most end of the concave upstream face to the longitudinal wall downstream of the inward-most end of the concave upstream face, wherein at least a portion of the concave upstream face extends downstream of the inward-most end of the concave upstream face, and wherein the concave upstream face provides a sheltered cavity at the rotating detonation combustion region at which a detonation wave propagates; a first fuel injector providing a first flow of fuel to the rotating detonation combustion region, and a second fuel injector providing a second flow of fuel through the downstream face to the additional downstream combustion region, which is positioned adjacent the downstream face and downstream of the rotating detonation combustion region; wherein the longitudinal wall further defines a combustion section at which the inner wall assembly is disposed and in which the rotation detonation combustion region and additional combustion region are located; and wherein the longitudinal wall further defines an exhaust nozzle distinct from and downstream of the combustion section in serial flow arrangement.
2. The engine of claim 1, further defining a first fuel injection port of the first fuel injector providing the first flow of fuel to the rotating detonation combustion region, and a second fuel injection port of the second fuel injector providing the second flow of fuel to the additional combustion region downstream of the rotating detonation combustion region.
3. The engine of claim 2, wherein the longitudinal wall defines the first fuel injection port providing the first flow of fuel to the rotating detonation combustion region.
4. The engine of claim 2, wherein the inner wall assembly defines the first fuel injection port through the concave upstream face providing the first flow of fuel to the rotating detonation combustion region.
5. The engine of claim 2, wherein the inner wall assembly defines the second fuel injection port through the downstream face providing the second flow of fuel for deflagrative combustion in the additional combustion region downstream of the rotating detonation combustion region.
6. The engine of claim 1, wherein the inward-most end of the concave upstream face defines a tip extended into the gas flowpath, and wherein the tip is further extended along a height, a width, or both, of the gas flowpath or an annulus of the gas flowpath.
7. The engine of claim 1, wherein the inner wall assembly is extended to approximately 35% or less of a depth of the gas flowpath.
8. The engine of claim 1, wherein the longitudinal wall defines a two dimensional gas flowpath defining a height and a width.
9. The engine of claim 1, wherein the longitudinal wall defines a first longitudinal wall and a second longitudinal wall each extended annularly around an axial centerline of the engine, and wherein the first longitudinal wall is outward radially of the second longitudinal wall.
10. The engine of claim 9, wherein the inner wall assembly is defined from the first longitudinal wall into the gas flowpath, wherein the inner wall assembly defines a first detonation combustion region adjacent to the first longitudinal wall.
11. The engine of claim 9, wherein the inner wall assembly is defined from the second longitudinal wall into the gas flowpath, wherein the inner wall assembly defines a second detonation combustion region adjacent to the second longitudinal wall.
12. The engine of claim 1, wherein the longitudinal wall further defines: an inlet section upstream of the combustion section in serial flow arrangement.
13. The engine of claim 12, wherein the inlet section is configured to admit a supersonic flow of oxidizer to the combustion section.
14. The engine of claim 12, wherein the longitudinal wall defines a contour at the inlet section to provide a subsonic flow of oxidizer to the combustion section.
15. The engine of claim 1, further comprising: a first fuel injection port of the first fuel injector configured to provide the first flow of fuel to the rotating detonation combustion region, wherein the first fuel injection port is positioned adjacent to the rotating detonation combustion region along the lengthwise direction; and a second fuel injection port of the second fuel injector wherein the second fuel injection port is configured to provide the second flow of fuel for a deflagrative combustion process downstream of the rotating detonation combustion region.
16. The engine of claim 1, further comprising an inlet upstream of the inner wall assembly, the inlet defining an upstream portion of the gas flowpath, wherein a first portion of a flow of oxidizer in the first portion of the gas flowpath is captured at the sheltered cavity for generating the rotating detonation wave, and a second portion of the flow of oxidizer through the inlet flows through the second portion of the gas flowpath to bypass the sheltered cavity and flow into the additional combustion region downstream of the rotating detonation region.
17. The engine of claim 1, wherein a portion of combustion gases from the rotating detonation wave flow into the second portion of the gas flowpath and into the additional combustion region downstream of the rotating detonation region.
Claims 21-40 are rejected on the ground of nonstatutory double patenting as being unpatentable over claims 1-17 of U.S. Patent No. 11959441 in view of the prior art applied above involving the CN ‘145 or the Morrison et al (3,240,010) combinations. The other prior art applied above are also applicable for the dependent claims and are incorporated from above for the various combinations involving Morrison or CN’145 and not repeated for conciseness. US 11959441 teaches the majority of the claimed limitations with minor variations in claim wording (e.g. exhaust section vs exhaust nozzle) and include an inlet section in claim 13 but do not clearly teach the first and second airflows are from the inlet section and contiguous to generate the respective first and second fuel/air mixtures and wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and the gas flowpath extending at least partially along a centerline of the combustion section [basically extends axially, wherein both the first portion and the second portion of the airflow are located within the gas flowpath. This is well known from the prior art of record applied above, particularly in light of Andreadis, to use the airflow from the inlet to supply the air from the air source to the gas flowpath which becomes the claimed first and second portions of airstream and wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is and contiguous and radially inward of the first portion of the airflow, the gas flowpath extending at least partially along a centerline of the combustion section [basically extends axially. It would have been obvious to one of ordinary skill in the art to employ an inlet to provide the claimed first and second portions of airstream to the first and second combustion regions, wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is contiguous with and radially inward of the first portion of the airflow and flowing along the central axis, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, the gas flowpath extending at least partially along a centerline of the combustion section [basically extends axially, as an obvious type of delivery system for the combustible gases needed to form first and second combustion regions. As for the limitation, wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, this is taught by e.g. applied prior art and incorporated by reference from the prior art treatment above. It would have been obvious to one of ordinary skill in the art to employ wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, in the manner applied above, as an obvious flow pattern used in the art for airflow. As for the supersonic engine, this is taught by CN ‘145 or Andreadis. This would be obvious as a typical use of rotating detonation and/or ramjets / scramjets. As for the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture, this is well known in the prior art Morrison does not teach specifically using a scramjet. Andreadis teaches the use of dual mode ramjets and scramjets [with subsonic and supersonic combustion – see paragraph 0002]. It would have been obvious to one of ordinary skill in the art to employ the combustor of Morrison in a scramjet, i.e. with supersonic combustion, as an Andreadis teaches using the same combustor structure in both ramjets and scramjets. The second combustion chamber added by any of Townend, Rich and Pavia, will be either an augmenter or ramjet / scramjet combustor. Note that Townend specifically teaches the second combustor is piloted by the first combustor, and Rich and Pavia do not teach an igniter is necessary and thus inherently piloted by the hot combustion gases, in the manner of Townend. Alternately, Andreadis teach the second fuel 108 is piloted by the first combustor 38 [e.g. paragraph 0026]. It would have been obvious to one of ordinary skill in the art to make the first combustor function as a pilot to ignite the second fuel-air mixture, as taught by either Townend or Andreadis, as the typical practice in the art that promotes simplicity of operation by making the second combustor ignitorless.
For claim 33, a change in preamble from a combustion section vs an engine with claimed combustion section, is an obvious variation from an engine with the claimed combustion section. It would have been obvious to employ the combustion section for a supersonic engine as a typical application of the claimed patent combustor.
18643376
US 11959441
A supersonic engine comprising;
an inlet section to which an inlet airflow is provided; a combustion section that receives an airflow from the inlet section, wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and flowing along the central axis; and an exhaust section [exhaust nozzle in ‘579 patent], a gas flowpath defined extending through the inlet section, the combustion section, and the exhaust section, the gas flowpath extending at least partially along a centerline of the combustion section, wherein both the first portion and the second portion of the airflow are located within the gas flowpath,
wherein the combustion section includes (a) a first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with a first portion of the airflow from the inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor, and
(b) a second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with a second portion of the airflow from the inlet section and with combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture.
22. The supersonic engine according to claim 21, wherein the supersonic engine is a ramjet engine and the second combustion portion is one of a ramjet combustor or an augmenter.
23. The supersonic engine according to claim 21, wherein the supersonic engine is a scramjet engine and the second combustion portion is one of a scramjet combustor or an augmenter.
24. The supersonic engine according to claim 21, wherein the combustion section includes a longitudinal wall extending along a length of the combustion section and defining an outer periphery of the first combustion portion and the second combustion portion, the gas flowpath being defined within the outer periphery of the combustion portion and the second combustion portion.
25. The supersonic engine according to claim 24, wherein the rotating detonation combustor extends about the outer periphery of the first combustion portion.
26. The supersonic engine according to claim 25, wherein the first combustion portion includes an inner wall assembly extending from the longitudinal wall into the gas flowpath, the inner wall assembly having an upstream face and a downstream wall.
27. The supersonic engine according to claim 26, wherein the inner wall assembly, downstream of the combustion chamber, extends about a periphery of the longitudinal wall and defines the gas flowpath therethrough.
28. The supersonic engine according to claim 26, wherein a rotating detonation region is defined at the upstream face of the inner wall assembly.
29. The supersonic engine according to claim 26, wherein the upstream face of the inner wall assembly captures the first portion of the airflow from the inlet section.
30. The supersonic engine according to claim 26, wherein the upstream face of the inner wall assembly includes an inner tip extending into the gas flowpath, and the downstream wall extends from the inner tip to the longitudinal wall.
31. The supersonic engine according to claim 30, wherein the combustion products from the rotating detonation combustor flow over the inner tip into the gas flowpath to mix with the second portion of the airflow.
32. The supersonic engine according to claim 30, wherein the downstream wall includes at least one second fuel injection port, and the second flow of fuel is provided into the second combustion portion via the at least one second fuel injection port.
33. A combustion section for a supersonic engine, comprising: a first combustion portion and a second combustion portion defining a gas flowpath therethrough, wherein the combustion section defines a central axis, wherein an airflow into the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is contiguous with and radially inward of the first portion of the airflow, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, the gas flowpath extending at least partially along a centerline of the combustion section, the first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with a first portion of an airflow received by the first combustion portion from an inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor; and the second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with a second portion of the airflow from the inlet section and with combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture.
1. An engine, comprising: a longitudinal wall extended along a lengthwise direction, wherein the longitudinal wall defines a gas flowpath of the engine; an inner wall assembly extended from the longitudinal wall into the gas flowpath, wherein the inner wall assembly defines a rotating detonation combustion region in a first portion of the gas flowpath upstream of the inner wall assembly and adjacent to the longitudinal wall along the lengthwise direction and extending inward a depth from the longitudinal wall, a second portion of the gas flowpath being surrounded by the rotating detonation combustion region and extending in the lengthwise direction from upstream of the rotating detonation combustion region to downstream of the rotating detonation combustion region into an additional combustion region downstream of the rotating detonation combustion region, wherein the inner wall assembly comprises (a) an upstream face extended from the longitudinal wall into the gas flowpath adjacent to the rotating detonation combustion region, and (b) a downstream face extending at least partially along the lengthwise direction from an inward-most end of the upstream face to the longitudinal wall downstream of the inward-most end of the upstream face, wherein at least a portion of the upstream face extends downstream of the inward-most end of the upstream face, and wherein the upstream face provides a sheltered cavity at the rotating detonation combustion region at which a detonation wave propagates; a first fuel injector providing a first flow of fuel to the rotating detonation combustion region, and a second fuel injector providing a second flow of fuel through the downstream face to the additional downstream combustion region, which is positioned adjacent the downstream face and downstream of the rotating detonation combustion region, wherein the longitudinal wall further defines a combustion section at which the inner wall assembly is disposed and in which the rotation detonation combustion region and additional combustion region are located, and wherein the longitudinal wall further defines an exhaust nozzle distinct from and downstream of the combustion section in serial flow arrangement; and an inlet section upstream of the combustion section in serial flow arrangement, wherein the longitudinal wall defines a contour at the inlet section to provide a subsonic flow of oxidizer to the rotating detonation combustion region.
2. The engine according to claim 1, further defining a first fuel injection port of the first fuel injector providing the first flow of fuel to the rotating detonation combustion region, and a second fuel injection port of the second fuel injector providing the second flow of fuel to the additional combustion region downstream of the rotating detonation combustion region.
3. The engine according to claim 2, wherein the longitudinal wall defines the first fuel injection port providing the first flow of fuel to the rotating detonation combustion region.
4. The engine according to claim 2, wherein the inner wall assembly defines the first fuel injection port through the upstream face providing the first flow of fuel to the rotating detonation combustion region.
5. The engine according to claim 2, wherein the inner wall assembly defines the second fuel injection port through the downstream face providing the second flow of fuel for deflagrative combustion in the additional combustion region downstream of the rotating detonation combustion region.
6. The engine according to claim 2, wherein the second fuel injector includes a third fuel injection port arranged through the downstream face downstream of the second fuel injection port.
7. The engine according to claim 1, wherein the inward-most end of the upstream face defines a tip extended into the gas flowpath, and wherein the tip is further extended along a height, a width, or both, of the gas flowpath or an annulus of the gas flowpath.
8. The engine according to claim 1, wherein the inner wall assembly is extended to approximately 35% or less of a depth of the gas flowpath.
9. The engine according to claim 1, wherein the longitudinal wall defines a two dimensional gas flowpath defining a height and a width.
10. The engine according to claim 1, wherein the longitudinal wall defines a first longitudinal wall and a second longitudinal wall each extended annularly around an axial centerline of the engine, and wherein the first longitudinal wall is outward radially of the second longitudinal wall.
11. The engine according to claim 10, wherein the inner wall assembly is defined from the first longitudinal wall into the gas flowpath, wherein the inner wall assembly defines a first detonation combustion region adjacent to the first longitudinal wall.
12. The engine according to claim 11, wherein the inner wall assembly is defined from the second longitudinal wall into the gas flowpath, wherein the inner wall assembly defines a second detonation combustion region adjacent to the second longitudinal wall.
13. The engine according to claim 1, wherein the longitudinal wall defines the exhaust nozzle to accelerate a flow of combustion gases from the combustion section.
14. The engine according to claim 1, wherein the inlet section includes a converging portion and a diverging portion in serial flow relationship.
15. The engine according to claim 1, wherein the upstream face is a concave upstream face extended from the longitudinal wall into the gas flowpath.
16. The engine according to claim 1, further comprising: a first fuel injection port of the first fuel injector configured to provide the first flow of fuel to the rotating detonation combustion region, wherein the first fuel injection port is positioned adjacent to the rotating detonation combustion region along the lengthwise direction; and a second fuel injection port of the second fuel injector, wherein the second fuel injection port is configured to provide the second flow of fuel for a deflagrative combustion process downstream of the rotating detonation combustion region.
17. The engine according to claim 1, wherein the inlet section defining an upstream portion of the gas flowpath, wherein a first portion of a flow of oxidizer in the first portion of the gas flowpath is captured at the sheltered cavity for generating a rotating detonation wave, and a second portion of the flow of oxidizer through the inlet section flows through the second portion of the gas flowpath to bypass the sheltered cavity and flow into the additional combustion region downstream of the rotating detonation combustion region.
18. The engine according to claim 17, wherein a portion of combustion gases from the rotating detonation wave flow into the second portion of the gas flowpath and into the additional combustion region downstream of the rotating detonation combustion region.
19. The engine according to claim 1, wherein the longitudinal wall further extends along a width and a height to define a rectangular cross section of the gas flowpath.
20. The engine according to claim 19, wherein the width and the height define a substantially square cross sectional area.
Claims 21-40 are rejected on the ground of nonstatutory double patenting as being unpatentable over claims 1-17 of U.S. Patent No. 11970994 in view of the prior art applied above involving the CN ‘145 or the Morrison et al (3,240,010) combinations. The other prior art applied above are also applicable for the dependent claims and are incorporated from above for the various combinations involving Morrison or CN’145 and not repeated for conciseness. US 11970994 teaches the majority of the claimed limitations with minor variations in claim wording (e.g. exhaust section vs exhaust nozzle) and include an inlet section in claim 13 but do not clearly teach the first and second airflows are from the inlet section and contiguous to generate the respective first and second fuel/air mixtures and wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and the gas flowpath extending at least partially along a centerline of the combustion section [basically extends axially, wherein both the first portion and the second portion of the airflow are located within the gas flowpath. This is well known from the prior art of record applied above, particularly in light of Andreadis, to use the airflow from the inlet to supply the air from the air source to the gas flowpath which becomes the claimed first and second portions of airstream and wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is and contiguous and radially inward of the first portion of the airflow, the gas flowpath extending at least partially along a centerline of the combustion section [basically extends axially. It would have been obvious to one of ordinary skill in the art to employ an inlet to provide the claimed first and second portions of airstream to the first and second combustion regions, wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is contiguous with and radially inward of the first portion of the airflow and flowing along the central axis, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, the gas flowpath extending at least partially along a centerline of the combustion section [basically extends axially, as an obvious type of delivery system for the combustible gases needed to form first and second combustion regions. As for the limitation, wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, this is taught by e.g. applied prior art and incorporated by reference from the prior art treatment above. It would have been obvious to one of ordinary skill in the art to employ wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, in the manner applied above, as an obvious flow pattern used in the art for airflow. As for the supersonic engine, this is taught by CN ‘145 or Andreadis. This would be obvious as a typical use of rotating detonation and/or ramjets / scramjets. As for the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture, this is well known in the prior art Morrison does not teach specifically using a scramjet. Andreadis teaches the use of dual mode ramjets and scramjets [with subsonic and supersonic combustion – see paragraph 0002]. It would have been obvious to one of ordinary skill in the art to employ the combustor of Morrison in a scramjet, i.e. with supersonic combustion, as an Andreadis teaches using the same combustor structure in both ramjets and scramjets. The second combustion chamber added by any of Townend, Rich and Pavia, will be either an augmenter or ramjet / scramjet combustor. Note that Townend specifically teaches the second combustor is piloted by the first combustor, and Rich and Pavia do not teach an igniter is necessary and thus inherently piloted by the hot combustion gases, in the manner of Townend. Alternately, Andreadis teach the second fuel 108 is piloted by the first combustor 38 [e.g. paragraph 0026]. It would have been obvious to one of ordinary skill in the art to make the first combustor function as a pilot to ignite the second fuel-air mixture, as taught by either Townend or Andreadis, as the typical practice in the art that promotes simplicity of operation by making the second combustor ignitorless.
For claim 33, a change in preamble from a combustion section vs an engine with claimed combustion section, is an obvious variation from an engine with the claimed combustion section. It would have been obvious to employ the combustion section for a supersonic engine as a typical application of the claimed patent combustor.
18643376
US 11970994
A supersonic engine comprising;
an inlet section to which an inlet airflow is provided; a combustion section that receives an airflow from the inlet section, wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and flowing along the central axis; and an exhaust section [exhaust nozzle in ‘579 patent], a gas flowpath defined extending through the inlet section, the combustion section, and the exhaust section, the gas flowpath extending at least partially along a centerline of the combustion section, wherein both the first portion and the second portion of the airflow are located within the gas flowpath,
wherein the combustion section includes (a) a first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with a first portion of the airflow from the inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor, and
(b) a second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with a second portion of the airflow from the inlet section and with combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture.
22. The supersonic engine according to claim 21, wherein the supersonic engine is a ramjet engine and the second combustion portion is one of a ramjet combustor or an augmenter.
23. The supersonic engine according to claim 21, wherein the supersonic engine is a scramjet engine and the second combustion portion is one of a scramjet combustor or an augmenter.
24. The supersonic engine according to claim 21, wherein the combustion section includes a longitudinal wall extending along a length of the combustion section and defining an outer periphery of the first combustion portion and the second combustion portion, the gas flowpath being defined within the outer periphery of the combustion portion and the second combustion portion.
25. The supersonic engine according to claim 24, wherein the rotating detonation combustor extends about the outer periphery of the first combustion portion.
26. The supersonic engine according to claim 25, wherein the first combustion portion includes an inner wall assembly extending from the longitudinal wall into the gas flowpath, the inner wall assembly having an upstream face and a downstream wall.
27. The supersonic engine according to claim 26, wherein the inner wall assembly, downstream of the combustion chamber, extends about a periphery of the longitudinal wall and defines the gas flowpath therethrough.
28. The supersonic engine according to claim 26, wherein a rotating detonation region is defined at the upstream face of the inner wall assembly.
29. The supersonic engine according to claim 26, wherein the upstream face of the inner wall assembly captures the first portion of the airflow from the inlet section.
30. The supersonic engine according to claim 26, wherein the upstream face of the inner wall assembly includes an inner tip extending into the gas flowpath, and the downstream wall extends from the inner tip to the longitudinal wall.
31. The supersonic engine according to claim 30, wherein the combustion products from the rotating detonation combustor flow over the inner tip into the gas flowpath to mix with the second portion of the airflow.
32. The supersonic engine according to claim 30, wherein the downstream wall includes at least one second fuel injection port, and the second flow of fuel is provided into the second combustion portion via the at least one second fuel injection port.
33. A combustion section for a supersonic engine, comprising: a first combustion portion and a second combustion portion defining a gas flowpath therethrough, wherein the combustion section defines a central axis, wherein an airflow into the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is contiguous with and radially inward of the first portion of the airflow , wherein both the first portion and the second portion of the airflow are located within the gas flowpath, the gas flowpath extending at least partially along a centerline of the combustion section, the first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with a first portion of an airflow received by the first combustion portion from an inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor; and the second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with a second portion of the airflow from the inlet section and with combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture.
1. An engine, comprising: a longitudinal wall extended along a lengthwise direction, wherein the longitudinal wall defines a gas flowpath of the engine; an inner wall assembly extended from the longitudinal wall into the gas flowpath,
wherein the inner wall assembly defines a rotating detonation combustion region in a first portion of the gas flowpath upstream of the inner wall assembly and adjacent to the longitudinal wall along the lengthwise direction and extending inward a depth from the longitudinal wall, a second portion of the gas flowpath being surrounded by the rotating detonation combustion region and extending in the lengthwise direction from upstream of the rotating detonation combustion region to downstream of the rotating detonation combustion region into an additional combustion region downstream of the rotating detonation combustion region, wherein the inner wall assembly comprises (a) an upstream face extended from the longitudinal wall into the gas flowpath adjacent to the rotating detonation combustion region, the upstream face being defined perpendicular from the longitudinal wall, and (b) a downstream face extending at least partially along the lengthwise direction from an inward-most end of the upstream face to the longitudinal wall downstream of the inward-most end of the upstream face, wherein the upstream face provides a sheltered cavity at the rotating detonation combustion region at which a detonation wave propagates; and a first fuel injector providing a first flow of fuel to the rotating detonation combustion region, and a second fuel injector providing a second flow of fuel through the downstream face to the additional downstream combustion region, which is positioned adjacent the downstream face and downstream of the rotating detonation combustion region, wherein the longitudinal wall further defines a combustion section at which the inner wall assembly is disposed and in which the rotation detonation combustion region and additional combustion region are located, and wherein the longitudinal wall further defines an exhaust nozzle distinct from and downstream of the combustion section in serial flow arrangement.
2. The engine according to claim 1, further defining a first fuel injection port of the first fuel injector providing the first flow of fuel to the rotating detonation combustion region, and a second fuel injection port of the second fuel injector providing the second flow of fuel to the additional combustion region downstream of the rotating detonation combustion region.
3. The engine according to claim 2, wherein the longitudinal wall defines the first fuel injection port providing the first flow of fuel to the rotating detonation combustion region.
4. The engine according to claim 2, wherein the inner wall assembly defines the first fuel injection port through the upstream face providing the first flow of fuel to the rotating detonation combustion region.
5. The engine according to claim 2, wherein the inner wall assembly defines the second fuel injection port through the downstream face providing the second flow of fuel for deflagrative combustion in the additional combustion region downstream of the rotating detonation combustion region.
6. The engine according to claim 2, wherein the second fuel injector includes a third fuel injection port arranged through the downstream face downstream of the second fuel injection port.
7. The engine according to claim 1, wherein the inward-most end of the upstream face defines a tip extended into the gas flowpath, and wherein the tip is further extended along a height, a width, or both, of the gas flowpath or an annulus of the gas flowpath.
8. The engine according to claim 1, wherein the inner wall assembly is extended to approximately 35% or less of a depth of the gas flowpath.
9. The engine according to claim 1, wherein the longitudinal wall defines a two dimensional gas flowpath defining a height and a width.
10. The engine according to claim 1, wherein the longitudinal wall defines a first longitudinal wall and a second longitudinal wall each extended annularly around an axial centerline of the engine, and wherein the first longitudinal wall is outward radially of the second longitudinal wall.
11. The engine according to claim 10, wherein the inner wall assembly is defined from the first longitudinal wall into the gas flowpath, wherein the inner wall assembly defines a first detonation combustion region adjacent to the first longitudinal wall.
12. The engine according to claim 11, wherein the inner wall assembly is defined from the second longitudinal wall into the gas flowpath, wherein the inner wall assembly defines a second detonation combustion region adjacent to the second longitudinal wall.
13. The engine according to claim 1, wherein the longitudinal wall further defines an inlet section upstream of the combustion section in serial flow arrangement.
14. The engine according to claim 13, wherein the inlet section is configured to admit a supersonic flow of oxidizer to the combustion section.
15. The engine according to claim 14, wherein the inlet section includes a converging inlet portion.
16. The engine according to claim 1, further comprising: a first fuel injection port of the first fuel injector configured to provide the first flow of fuel to the rotating detonation combustion region, wherein the first fuel injection port is positioned adjacent to the rotating detonation combustion region along the lengthwise direction; and a second fuel injection port of the second fuel injector, wherein the second fuel injection port is configured to provide the second flow of fuel for a deflagrative combustion process downstream of the rotating detonation combustion region.
17. The engine according to claim 1, further comprising an inlet upstream of the inner wall assembly, the inlet defining an upstream portion of the gas flowpath, wherein a first portion of a flow of oxidizer in the first portion of the gas flowpath is captured at the sheltered cavity for generating a rotating detonation wave, and a second portion of the flow of oxidizer through the inlet flows through the second portion of the gas flowpath to bypass the sheltered cavity and flow into the additional combustion region downstream of the rotating detonation combustion region.
18. The engine according to claim 17, wherein a portion of combustion gases from the rotating detonation wave flow into the second portion of the gas flowpath and into the additional combustion region downstream of the rotating detonation combustion region.
19. The engine according to claim 1, wherein the longitudinal wall further extends along a width and a height to define a rectangular cross section of the gas flowpath.
20. The engine according to claim 19, wherein the width and the height define a substantially square cross sectional area.
Claims 21-40 are rejected on the ground of nonstatutory double patenting as being unpatentable over claims 1-9 of U.S. Patent No. 11320147 in view of the prior art applied above involving the CN ‘145 or the Morrison et al (3,240,010) combinations. The other prior art applied above are also applicable for the dependent claims and are incorporated from above for the various combinations involving Morrison or CN’145 and not repeated for conciseness. US 11320147 teaches the majority of the claimed limitations and do not teach an exhaust section nor an inlet section in claim 13 and do not clearly teach the first and second airflows are from the inlet section and contiguous to generate the respective first and second fuel/air mixtures and wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is radially inward of the first portion of the airflow and the gas flowpath extending at least partially along a centerline of the combustion section [basically extends axially, wherein both the first portion and the second portion of the airflow are located within the gas flowpath. This is well known from the prior art of record applied above, particularly in light of Andreadis, to use the airflow from the inlet to supply the air from the air source to the gas flowpath which becomes the claimed first and second portions of airstream and wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is and contiguous and radially inward of the first portion of the airflow, the gas flowpath extending at least partially along a centerline of the combustion section [basically extends axially. As for the limitation, wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, this is taught by e.g. applied prior art and incorporated by reference from the prior art treatment above. It would have been obvious to one of ordinary skill in the art to employ wherein the airflow received by the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is radially contiguous with and radially in between a first radially outer stream of the first portion of the airflow and a radially inner stream of the first portion of the airflow, in the manner applied above, as an obvious flow pattern used in the art for airflow. It would have been obvious to one of ordinary skill in the art to employ an inlet to provide the claimed first and second portions of airstream to the first and second combustion regions, wherein the combustion section defines a central axis, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is contiguous with and radially inward of the first portion of the airflow and flowing along the central axis, wherein both the first portion and the second portion of the airflow are located within the gas flowpath, the gas flowpath extending at least partially along a centerline of the combustion section [basically extends axially, as an obvious type of delivery system for the combustible gases needed to form first and second combustion regions. It would have been obvious to utilize an exhaust section for the combustion gases to escape. As for the supersonic engine, this is taught by CN ‘145 or Andreadis. This would be obvious as a typical use of rotating detonation and/or ramjets / scramjets. As for the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture, this is well known in the prior art Morrison does not teach specifically using a scramjet. Andreadis teaches the use of dual mode ramjets and scramjets [with subsonic and supersonic combustion – see paragraph 0002]. It would have been obvious to one of ordinary skill in the art to employ the combustor of Morrison in a scramjet, i.e. with supersonic combustion, as an Andreadis teaches using the same combustor structure in both ramjets and scramjets. The second combustion chamber added by any of Townend, Rich and Pavia, will be either an augmenter or ramjet / scramjet combustor. Note that Townend specifically teaches the second combustor is piloted by the first combustor, and Rich and Pavia do not teach an igniter is necessary and thus inherently piloted by the hot combustion gases, in the manner of Townend. Alternately, Andreadis teach the second fuel 108 is piloted by the first combustor 38 [e.g. paragraph 0026]. It would have been obvious to one of ordinary skill in the art to make the first combustor function as a pilot to ignite the second fuel-air mixture, as taught by either Townend or Andreadis, as the typical practice in the art that promotes simplicity of operation by making the second combustor ignitorless.
The patent does not necessarily teach the gas flowpath being an open flowpath excluding a centerbody therewithin [claim 21, 27, 37]; however, this is regarded as inherent as there was no centerbody claimed. Alternately, Morrison would teach that eliminating a centerbody [excluded in larger engines, col. 6, lines 44-59] / open flow path is well known for rotating detonation combustors and it would have been obvious to do so to reduce flow losses or for simplicity of construction. For claim 33, a change in preamble from a combustion section vs an engine with claimed combustion section, is an obvious variation from an engine with the claimed combustion section. It would have been obvious to employ the combustion section for a supersonic engine as a typical application of the claimed patent combustor.
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17935646
US 11320147
21. A supersonic engine comprising;
an inlet section to which an inlet airflow is provided; a combustion section that receives an airflow from the inlet section, wherein the airflow includes a first portion and a second portion, and wherein the second portion of the airflow is contiguous with and radially inward of the first portion of the airflow; and an exhaust section [exhaust nozzle in ‘579 patent], a gas flowpath defined extending through the inlet section, the combustion section, and the exhaust section, the gas flowpath extending at least partially along a centerline of the combustion section,, wherein both the first portion and the second portion of the airflow are located within the open flowpath,
wherein the combustion section includes (a) a first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with a first portion of the airflow from the inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor, and
(b) a second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with a second portion of the airflow from the inlet section and with combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture.
22. The supersonic engine according to claim 21, wherein the supersonic engine is a ramjet engine and the second combustion portion is one of a ramjet combustor or an augmenter.
23. The supersonic engine according to claim 21, wherein the supersonic engine is a scramjet engine and the second combustion portion is one of a scramjet combustor or an augmenter.
24. The supersonic engine according to claim 21, wherein the combustion section includes a longitudinal wall extending along a length of the combustion section and defining an outer periphery of the first combustion portion and the second combustion portion, the gas flowpath being defined within the outer periphery of the combustion portion and the second combustion portion.
25. The supersonic engine according to claim 24, wherein the rotating detonation combustor extends about the outer periphery of the first combustion portion.
26. The supersonic engine according to claim 25, wherein the first combustion portion includes an inner wall assembly extending from the longitudinal wall into the gas flowpath, the inner wall assembly having an upstream face and a downstream wall.
27. The supersonic engine according to claim 26, wherein the inner wall assembly, downstream of the combustion chamber, extends about a periphery of the longitudinal wall and defines the gas flowpath therethrough.
28. The supersonic engine according to claim 26, wherein a rotating detonation region is defined at the upstream face of the inner wall assembly.
29. The supersonic engine according to claim 26, wherein the upstream face of the inner wall assembly captures the first portion of the airflow from the inlet section.
30. The supersonic engine according to claim 26, wherein the upstream face of the inner wall assembly includes an inner tip extending into the gas flowpath, and the downstream wall extends from the inner tip to the longitudinal wall.
31. The supersonic engine according to claim 30, wherein the combustion products from the rotating detonation combustor flow over the inner tip into the gas flowpath to mix with the second portion of the airflow.
32. The supersonic engine according to claim 30, wherein the downstream wall includes at least one second fuel injection port, and the second flow of fuel is provided into the second combustion portion via the at least one second fuel injection port.
33. A combustion section for a supersonic engine, comprising: a first combustion portion and a second combustion portion defining a gas flowpath therethrough, wherein the combustion section defines a central axis, wherein an airflow into the combustion section includes a first portion and a second portion, and wherein the second portion of the airflow is contiguous with and radially inward of the first portion of the airflow , wherein both the first portion and the second portion of the airflow are located within the gas flowpath, the gas flowpath extending at least partially along a centerline of the combustion section, the first combustion portion having a rotating detonation combustor to which a first flow of fuel is provided to mix with a first portion of an airflow received by the first combustion portion from an inlet section to generate a first fuel-air mixture in the first combustion portion that is detonated to generate rotating detonation waves within the rotating detonation combustor; and the second combustion portion arranged downstream of the rotating detonation combustor, a second flow of fuel being provided to the second combustion portion to mix with a second portion of the airflow from the inlet section and with combustion products from the rotating detonation combustor to generate a second fuel-air mixture, the combustion products from the rotating detonation combustor functioning as a pilot to ignite the second fuel-air mixture.
1. An engine, the engine comprising: a longitudinal wall extended along a lengthwise direction, wherein the longitudinal wall defines a gas flowpath and a combustion section; an inner wall assembly extendable from the longitudinal wall into the gas flowpath, wherein the inner wall assembly includes an upstream face and defines an adjustable rotating detonation combustion region at the upstream face in the gas flowpath adjacent to the longitudinal wall, and wherein the inner wall assembly is extendable into the gas flowpath to adjust a depth of the rotating detonation combustion region in correspondence with an adjustable depth of the upstream face; an actuator coupled to the inner wall assembly, wherein the actuator is configured to actuate the inner wall assembly along a depth of the gas flowpath perpendicular to a direction of an oxidizer flowing through the combustion section to adjust the depth of the adjustable rotating detonation combustion region via adjusting the depth of the upstream face of the inner wall assembly; and wherein the engine is configured to perform operations comprising; flowing the oxidizer through the gas flowpath into the combustion section, wherein the flow of the oxidizer includes a first portion and a second portion; capturing, via the inner wall assembly being extended the depth into the gas flowpath by the actuator, the first portion of the flow of the oxidizer at the adjustable rotating detonation combustion region; flowing a first flow of fuel to the first portion of the flow of the oxidizer captured at the adjustable rotating detonation combustion region via the inner wall assembly; producing, via a first mixture of the first flow of fuel and the first portion of the oxidizer at the adjustable rotating detonation combustion region, a rotating detonation wave of detonation gases; and burning a second mixture comprising at least a portion of the detonation gases, a second flow of fuel, and the second portion of the flow of the oxidizer, wherein burning the second mixture is downstream of the rotating detonation wave of detonation gases.
2. The engine of claim 1, the operations comprising: adjusting, via the actuator, the inner wall assembly to between 0% and 35% of the depth of the gas flowpath.
3. The engine of claim 1, wherein the inner wall assembly comprises: a downstream face extended from the longitudinal wall and coupled to the upstream face, wherein the downstream face is disposed at an angle relative to the longitudinal wall.
4. The engine of claim 3, wherein the longitudinal wall or the upstream face of the inner wall assembly comprises a first fuel injection port, the first fuel injection port configured to provide the first flow of fuel to the adjustable rotating detonation combustion region.
5. The engine of claim 3, wherein the downstream face of the inner wall assembly comprises a second fuel injection port, the second fuel injection port configured to provide the second flow of fuel downstream of the adjustable rotating detonation combustion region relative to the flow of oxidizer through the gas flowpath.
6. The engine of claim 3, the operations comprising: adjusting an angle of the downstream face relative to the longitudinal wall between 10 degrees and 80 degrees relative to the longitudinal wall.
7. The engine of claim 1, the operations comprising: adjusting a cross-sectional area of the gas flowpath based on an operating condition of the engine.
8. The engine of claim 7, wherein adjusting the cross-sectional area of the gas flowpath comprises: adjusting, via the actuator, the depth of the inner wall assembly along the depth of the gas flowpath based at least on the operating condition of the engine.
9. The engine of claim 1, wherein the combustion section comprises: a deflagrative combustion process downstream of the detonation combustion region relative to the flow of oxidizer in the gas flowpath.
Response to Arguments
Applicant's arguments filed 12/16/2025 have been fully considered but they are not persuasive. Applicant’s arguments essentially allege the prior art do not teach the new limitations added by amendment. These limitations 1) raise issues of possession and indefiniteness and 2) are fully treated and above and are taught by the art of record. In response to applicant’s allegations regarding elements 1 and 2 in CN ‘145, these are not persuasive, as they are present in the annotated Fig. To address applicant’s concerns, it is noted that more annotations have been added to clarify the first and second portions of airflow. Applicant’s arguments concerning the combination with CN’348 are only a general allegation that one of ordinary skill in the art would not combined these and are not persuasive.
Applicant’s arguments concerning Morrison are similarly deficient as applicant did not possess nor clearly claim the invention. In response to applicant’s argument that there is no teaching, suggestion, or motivation to combine the references, the examiner recognizes that obviousness may be established by combining or modifying the teachings of the prior art to produce the claimed invention where there is some teaching, suggestion, or motivation to do so found either in the references themselves or in the knowledge generally available to one of ordinary skill in the art. See In re Fine, 837 F.2d 1071, 5 USPQ2d 1596 (Fed. Cir. 1988), In re Jones, 958 F.2d 347, 21 USPQ2d 1941 (Fed. Cir. 1992), and KSR International Co. v. Teleflex, Inc., 550 U.S. 398, 82 USPQ2d 1385 (2007). In this case, adding an additional second combustion portion is well within the ordinary skill in the art, as taught by the modifying references any of Townend (3,603,094), Rich (2,952,123), Pavia et al (2011/0005193) and further in view of Andreadis (2008/0196414). There are express motivating reasons set forth above, including “as taught by any of Townend, Rich and Pavia et al, in order to reduce noise and increase thrust as well as reduce drag at the exit of the nozzle and/or to increase the thrust of the jet engine or permit used for thrust vectoring and/or to cool the nozzle while reducing performance losses” Accordingly, applicant’s arguments fail to persuade.
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.
Contact Information
Any inquiry concerning this communication or earlier communications from the Examiner should be directed to TED KIM whose telephone number is 571-272-4829. The Examiner can be reached on regular business hours before 5:00 pm, Monday to Thursday and every other Friday.
The fax number for the organization where this application is assigned is 571-273-8300.
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/Ted Kim/
Telephone
571-272-4829
Primary Examiner
Fax
571-273-8300
March 26, 2026