CN115199438A - Turbine rotating rocket combined engine - Google Patents
Turbine rotating rocket combined engine Download PDFInfo
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- CN115199438A CN115199438A CN202210759257.XA CN202210759257A CN115199438A CN 115199438 A CN115199438 A CN 115199438A CN 202210759257 A CN202210759257 A CN 202210759257A CN 115199438 A CN115199438 A CN 115199438A
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- flow channel
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- oxidant
- combustion chamber
- engine
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- 238000002485 combustion reaction Methods 0.000 claims abstract description 52
- 239000000446 fuel Substances 0.000 claims abstract description 45
- 239000007789 gas Substances 0.000 claims abstract description 26
- 239000002737 fuel gas Substances 0.000 claims abstract description 19
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 claims abstract description 7
- 239000001301 oxygen Substances 0.000 claims abstract description 7
- 229910052760 oxygen Inorganic materials 0.000 claims abstract description 7
- 239000007800 oxidant agent Substances 0.000 claims description 66
- 230000001590 oxidative effect Effects 0.000 claims description 65
- 239000000295 fuel oil Substances 0.000 claims description 30
- 238000007789 sealing Methods 0.000 claims description 20
- 239000007921 spray Substances 0.000 claims description 10
- 239000003921 oil Substances 0.000 claims description 8
- 239000007787 solid Substances 0.000 claims description 6
- 239000003380 propellant Substances 0.000 abstract description 6
- 230000002269 spontaneous effect Effects 0.000 abstract description 2
- 230000010354 integration Effects 0.000 abstract 1
- 239000007788 liquid Substances 0.000 abstract 1
- 230000007246 mechanism Effects 0.000 description 8
- 230000003044 adaptive effect Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 238000010276 construction Methods 0.000 description 1
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
- 238000005859 coupling reaction Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000011160 research Methods 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
- F02K9/66—Combustion or thrust chambers of the rotary type
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention discloses a turbine rotary rocket combined engine, which comprises a turbine engine part and a rotary rocket part which are axially connected; the rotary rocket part removes a turbine pump, a combustion chamber and a nozzle of a traditional liquid fuel engine, a propellant is taken as a working medium, the propellant firstly passes through an oil delivery shaft before entering the rotary fuel gas generator, the oil delivery shaft is arranged on the axis of the engine and synchronously rotates with the engine, the propellant can be ignited by spontaneous combustion once contacting and colliding, and the integration of the fuel gas generator and shaft power output under the condition of no turbine structure can be realized without an additional ignition system; meanwhile, a afterburning chamber is added, so that the oxygen-enriched high-temperature gas generated by the turbine engine part and the oil-enriched high-temperature gas generated by the gas generator of the rotary rocket part are fully combusted in the afterburning chamber and then are ejected out through a thrust nozzle to generate thrust. The invention gives consideration to the overall performance of the engine under the conditions of high-speed and low-speed flight, widens the flight speed range and improves the thermodynamic cycle efficiency of the engine.
Description
Technical Field
The invention relates to the field of aerospace propulsion, in particular to a turbine rotary rocket combined engine.
Background
In recent years, with the progress and development of scientific and technical military, the adjacent space gradually draws attention from various countries due to the potential military and civil values of the adjacent space, and the countries carry out a series of researches on adjacent space aircrafts. The near space aircraft propulsion system is required to have the capability of accelerating and cruising at a high speed, and meanwhile, the near space aircraft propulsion system has the characteristics of high performance, light weight, high reliability, low cost, long voyage, reusability, wide range of adaptive speed and the like. The existing propulsion systems, including various aero-engines, ramjets, rocket engines, and piston engines, all have their own performance advantages and ideal flight airspace. The aircraft engine is characterized by high specific impulse performance, and is difficult to be used for a propulsion system of an aircraft with the flight speed greater than Mach 3.0 due to the limitation of temperature resistance limit of a turbine; the ramjet has higher performance and flight Mach number, but does not have self-acceleration capability, and the starting speed problem needs to be solved by means of a booster; the rocket engine is not limited by altitude and initial speed, but has low specific impulse performance and high propellant consumption rate.
By adopting a mode of different power combination cycles, for example, an air-breathing propulsion system is used under the condition of low flight Mach number, and the rocket engine is switched to work when the flight Mach number is higher, so that the maneuverability, flexibility and economy of the aircraft can be certainly improved. Therefore, the development of the air suction type combined power is significant. However, combined engines currently under development, involving turbine engines, such as turbine-based combined cycle engines and air turbine rocket engines, are limited by the temperature tolerance limits of the turbines. The temperature resistance limit of the turbine is always a core restriction factor for improving the thermodynamic cycle efficiency of the engine, and the influence of the temperature limit of the turbine on the thermodynamic cycle efficiency is more critical along with the increase of the temperature of the incoming flow under high-speed flight, and even becomes the key for determining whether the thermodynamic cycle mode is feasible or not.
Disclosure of Invention
Aiming at the problems in the prior art, the invention aims to provide a turbine rotary rocket combined engine, which has the following specific technical scheme:
a turbo-rocket engine assembly comprising a turbo-engine part and a rocket-rotating part axially connected;
the turbine engine part comprises an air inlet volute, an air compressor, a rotor shaft, a flame tube and a gas turbine;
the rotary rocket part comprises a lateral fuel oil sealing cavity, a lateral oxidant sealing cavity, a front bearing seat, a bearing sleeve, a combustion chamber body, a combustion chamber jet pipe, a afterburning chamber casing and a thrust jet pipe;
the afterburning chamber casing is fixedly connected to the air inlet volute; the thrust spray pipe is connected with the afterburning chamber casing; the combustion generator consisting of a lateral fuel oil sealing cavity, a lateral oxidant sealing cavity, a front bearing seat, a bearing sleeve, a combustion chamber body and a combustion chamber spray pipe is positioned in a afterburning chamber formed by the afterburning chamber casing;
the combustion chamber body comprises a main shaft extending out and a combustion chamber cavity arranged in the main shaft; the main shaft is a stepped concentric shaft sleeve shaft and comprises an inner solid section, a middle thin-walled section and an outer thin-walled section which are sleeved together; a first fuel flow passage is formed in the middle layer thin-wall section, a first oxidant flow passage is formed in the outer layer thin-wall section, and the first fuel flow passage and the first oxidant flow passage are both annular; the second fuel flow channels are uniformly distributed along the radial direction of the tail end of the first fuel flow channel, and the first fuel flow channels and the second fuel flow channels form fuel flow channels; the oxidant flow channels II are uniformly distributed along the radial direction of the tail end of the oxidant flow channel I, and the oxidant flow channel I and the oxidant flow channel II form an oxidant flow channel; the combustion chamber cavity is positioned between the plane where the second fuel flow channel is distributed and the plane where the second oxidant flow channel is distributed, one end of the second fuel flow channel is communicated with the first fuel flow channel, and the other end of the second fuel flow channel is communicated with the combustion chamber cavity through a fuel nozzle; one end of the oxidant flow channel II is communicated with the oxidant flow channel I, and the other end of the oxidant flow channel II is communicated with the combustion chamber cavity through an oxidant nozzle; a plurality of inclined exhaust outlets are uniformly formed in the rear end face of the combustion chamber cavity, and a combustion chamber spray pipe is arranged in each exhaust outlet;
the lateral fuel oil sealing cavity, the lateral oxidant sealing cavity, the front bearing seat and the bearing sleeve are sequentially and fixedly connected along the axis direction of the main shaft and sleeved on the main shaft; the lateral fuel oil sealing cavity is positioned between the inner solid section and the middle thin-walled section of the main shaft, and is provided with a fuel oil inlet along the radial direction;
and after the oxygen-rich high-temperature fuel gas generated by the turbine engine part and the oil-rich high-temperature fuel gas generated by the fuel gas generator of the rotary rocket part are combusted again in the afterburning chamber, the fuel gas is sprayed out through the thrust spray pipe to generate thrust.
Furthermore, axial oil grooves are uniformly distributed along the circumferential direction on the inner surfaces of the middle-layer thin-wall section and the outer-layer thin-wall section of the main shaft.
Further, the diameter of the section of the inner cavity of the combustion chamber nozzle is firstly reduced and then increased.
Further, the number of the spray pipes is 3.
Furthermore, the second fuel oil flow passage and the second oxidant flow passage are both six.
The invention has the following beneficial effects:
the turbine rotary rocket combined engine of the invention exerts the advantages of the turbine engine and the rotary rocket in the aspect of specific impulse and thrust-weight ratio as much as possible, can effectively solve the problem of serious restriction of turbine temperature limit on thermodynamic cycle efficiency, broadens the flight speed range, improves the working temperature of a gas generator and further improves the thermodynamic cycle efficiency of the engine so as to give consideration to the overall performance of the engine under high-speed and low-speed flight conditions, can effectively reduce the complexity of an engine system and improve the working reliability of the engine system.
Under the condition of low-speed flight, a turbine engine independent working mode with high specific impulse performance is adopted, and under the condition of high-speed flight, a turbine engine and rotary rocket engine combined working mode or an independent rotary rocket working mode is adopted, so that the high specific impulse performance of a low-speed stage can be kept, and the range of flight Mach number can be widened. Compared with the conventional air turbine rocket engine, the high-speed air turbine rocket engine can keep the working capacity under high Mach number and can also keep the high specific impact performance under low Mach number, thereby improving the flight time and the flight distance. The invention combines the advantages of a turbine engine and a rocket engine, can take off horizontally, mainly adopts a turbine engine working mode and a turbine engine and rotary rocket common working mode in the atmosphere, and fully utilizes air as an oxidant; outside the atmosphere, the working mode of the rotary rocket engine is directly adopted for orbit entering, and free switching in and out in the adjacent space is realized.
Drawings
FIG. 1 is a profile view of a turbine rotary rocket assembly engine according to an exemplary embodiment;
FIG. 2 is a cross-sectional view of a turbine-rotary rocket assembly engine shown in accordance with an exemplary embodiment;
FIG. 3 is a schematic view of a rotary gasifier of the present invention;
FIG. 4 is a front sectional view of the rotary gas generator of the present invention;
fig. 5 is a schematic view of the spindle 501.
Fig. 6 isbase:Sub>A sectional view taken along linebase:Sub>A-base:Sub>A in fig. 4.
Fig. 7 is a cross-sectional view taken along line C-C in fig. 4.
Fig. 8 is a sectional view taken along line B-B in fig. 4.
Fig. 9 is a sectional view of the spout 7.
Fig. 10 is a cross-sectional view of the main shaft 501.
In the figure, a lateral fuel oil seal cavity 1, a lateral oxidant seal cavity 2, a front bearing seat 3, a bearing sleeve 4, a combustion chamber body 5 and a combustion chamber jet pipe 6 are arranged; the gas turbine comprises an air inlet volute 7, a gas compressor 8, a rotor shaft 9, a flame tube 10, a gas guider 11, a gas turbine 12 and a coupling 13; the afterburner comprises an afterburning chamber casing 14, a thrust jet pipe 15, a fuel inlet 101, an oxidant inlet 201, a main shaft 501, a fuel flow channel I502, a fuel flow channel II 503, an oxidant flow channel I504, an oxidant flow channel II 505, a fuel nozzle 506, an oxidant nozzle 507, a combustion chamber cavity 508, an exhaust outlet 509 and an oil groove 510.
Detailed Description
The present invention will be described in detail below with reference to the accompanying drawings and preferred embodiments, and the objects and effects of the present invention will become more apparent, it being understood that the specific embodiments described herein are merely illustrative of the present invention and are not intended to limit the present invention.
As shown in fig. 1 and 2, the turbo-rocket motor assembly of the present invention includes a turbine motor part and a rocket motor part.
The turbine engine part comprises an air inlet volute 7, a compressor 8, a rotor shaft 9, a flame tube 10, a gas guider 11 and a gas turbine 12. The air inlet volute 7 is positioned at the forefront of the whole combined engine and is used for guiding the airflow entering the compressor 8. The compressor 8, the liner 10, the gas guide 11 and the gas turbine 12 are arranged in sequence along the axial direction of the rotor shaft 9. The turbine engine of the present invention is of a construction well known to those skilled in the art and will not be described in detail herein.
The rotary rocket part comprises a lateral fuel oil seal cavity 1, a lateral oxidant seal cavity 2, a front bearing seat 3, a bearing sleeve 4, a combustion chamber body 5, a combustion chamber jet pipe 6, a afterburning chamber casing 14 and a thrust jet pipe 15.
The afterburning chamber casing 14 is fixedly connected with the thrust jet pipe 15, and a fuel gas generator consisting of a lateral fuel oil sealing cavity 1, a lateral oxidant sealing cavity 2, a front bearing seat 3, a bearing sleeve 4, a combustion chamber body 5 and a combustion chamber jet pipe 6 is positioned in an afterburning chamber in the afterburning chamber casing 14.
As shown in fig. 3 to 4, the combustion chamber body 5 includes a protruding main shaft 501 and an internal combustion chamber cavity 508. The main shaft 501 is connected with the rotor shaft 9 through a coupler, so that the two shafts rotate synchronously. The main shaft 501 is a stepped concentric shaft sleeve shaft, and as shown in fig. 4, includes an inner solid section, a middle thin-walled section, and an outer thin-walled section, which are sleeved together. As shown in FIG. 2, a first fuel flow passage 502 is formed in the middle thin-wall section, a first oxidant flow passage 504 is formed in the outer thin-wall section, and the first fuel flow passage 502 and the first oxidant flow passage 504 are both annular. The second fuel flow channel 503 is uniformly distributed along the radial direction of the tail end of the first fuel flow channel 502, and the two fuel flow channels form a fuel flow channel; the second oxidant flow channels 505 are uniformly distributed along the radial direction of the end of the first oxidant flow channel 504, and the two oxidant flow channels form an oxidant flow channel. The combustion chamber cavity 508 is positioned between the plane on which the second fuel flow channel 503 is distributed and the plane on which the second oxidant flow channel 505 is distributed, one end of the second fuel flow channel 503 is communicated with the first fuel flow channel 502, and the other end of the second fuel flow channel is communicated with the combustion chamber cavity 508 through the fuel nozzle 506; the second oxidant flow channel 505 has one end communicating with the first oxidant flow channel 504 and the other end communicating with the combustion chamber 508 through the oxidant nozzle 507. As shown in fig. 3 and 6, the rear end surface of the combustion chamber 508 is further opened with three inclined exhaust outlets 509, and each exhaust outlet is provided with a combustion chamber nozzle 6.
As shown in fig. 7 and 8, in this embodiment, six fuel flow passages 503 and six oxidant flow passages 505 are provided, and 3 combustion chamber nozzles 6 are provided. The specific number of the components can be set according to actual needs, and the components are guaranteed to be uniformly distributed. As shown in fig. 9, the diameter of the inner cavity of the combustion chamber nozzle 6 is reduced first and then increased, so that the velocity of the high-temperature gas passing through the throat is local sonic velocity, and the velocity of the high-temperature gas passing through the throat is supersonic velocity, and sufficient tangential thrust is generated to drive the compressor to rotate and provide sufficient pressure for the fuel.
The lateral fuel oil sealing cavity 1, the lateral oxidant sealing cavity 2, the front bearing seat 3 and the bearing sleeve 4 are sleeved on the main shaft 501 sequentially through the bearing along the axis direction of the main shaft 501. The flange plates of the lateral fuel oil seal cavity 1 and the lateral oxidant seal cavity 2 are fixedly connected together through bolts, the interiors of the front bearing seat 4 and the bearing sleeve 5 are also arranged on the outer thin-wall section of the main shaft 501 through bearings, and the lateral oxidant seal cavity 2, the front bearing seat 3 and the bearing sleeve 4 are fixedly connected through bolts.
The lateral fuel oil sealing cavity 1 is located between the inner solid section and the middle thin-walled section of the main shaft 501, the lateral fuel oil sealing cavity 1 is provided with a radial fuel oil inlet 101, and fuel oil enters the lateral fuel oil sealing cavity 1 from the fuel oil inlet 101 and then enters the combustion chamber cavity 508 through the fuel oil flow passage I502, the fuel oil flow passage II 503 and the fuel oil nozzle 506 in sequence. The lateral oxidant sealed cavity 2 is located between the middle layer thin-wall section and the outer layer thin-wall section of the main shaft 501, the lateral oxidant sealed cavity 2 is provided with a radial oxidant inlet 201, and oxidant enters from the oxidant inlet 201 and then sequentially enters the combustion chamber cavity 508 through the oxidant runner I504, the oxidant runner II 505 and the oxidant nozzle 507.
The working principle of the turbo-rocket engine assembly according to the present invention will be described.
For the turbine engine part, air is pressurized by an inlet flow channel of an air inlet volute 7 through an air compressor 8 and then enters a flame tube 10 from the air compressor flow channel of the air compressor 8, the air is combusted with fuel oil in a combustion area of the flame tube 10 to generate oxygen-rich high-temperature fuel gas, the oxygen-rich high-temperature fuel gas impacts a gas turbine 12 through a flow channel of a fuel gas guider 11 to drive the gas turbine 12 to rotate, and then the oxygen-rich high-temperature fuel gas enters a afterburning chamber casing 14 through a blade flow channel of the gas turbine 12;
for the rotary rocket part, fuel oil and oxidant are sprayed out from a nozzle through a fuel oil flow channel and an oxidant flow channel of the main shaft 501 respectively and then collide, atomize and ignite spontaneously, the fuel oil and the oxidant are fully combusted in the combustion chamber cavity 508 to generate oil-rich high-temperature fuel gas, the high-temperature fuel gas is accelerated through the combustion chamber spray pipe 6 which is obliquely arranged to generate tangential thrust, fuel gas exhaust and the main shaft 501 generate a certain angle a, and due to the angle, a tangential thrust component is generated to drive the main shaft 501 to rotate, so that the rotor shaft 9 fixedly connected with the main shaft 501 is driven to rotate synchronously.
For the whole turbine rotary rocket combined engine, oxygen-rich high-temperature gas generated by the turbine engine part and oil-rich high-temperature gas generated by the rotary rocket part are efficiently mixed in a afterburning chamber formed by an afterburning chamber casing 14 and are combusted again, and the completely combusted gas is ejected out through a thrust nozzle 15 to generate thrust.
The turbine rotary rocket combination engine can provide three working mechanisms, namely an independent turbine engine working mechanism, an independent rotary rocket working mechanism, a turbine engine and a rotary rocket composite working mechanism.
The turbine engine operating regime alone is suitable for low speed cruise tasks. Under the mechanism, air in the atmosphere is fully utilized as an oxidant, the combined power system is mainly started by a turbine engine, the compressor is directly driven to work, the starting can be completed in a very short time, the high specific impulse characteristic and the fuel economy are realized, and the voyage can be greatly improved.
The single rotary rocket working mechanism is suitable for the on-orbit operation of the outer space. Under the mechanism, shaft work is mainly generated by the rotary rocket to directly drive the gas compressor to work, and gas sucked by the gas compressor is combusted by high-temperature fuel gas in the afterburning chamber and the rotary rocket to generate thrust. The working capacity under higher Mach number can be realized, and the range of the flight Mach number can be widened.
The combined working mechanism of the turbine engine and the rotary rocket can fully take the advantages of the turbine engine and the rotary rocket engine, can keep high specific impulse performance in a low-speed stage, and can widen the range of flight Mach number. The adjustment of different thrust levels can be realized, and the optimal specific impulse and fuel economy can be optimally realized under different flight Mach numbers.
The turbine rotary rocket combined engine provided by the invention has three combustion chambers, namely a turbine engine combustion chamber, a rotary rocket combustion chamber and a afterburning chamber, and the freedom degree of design and control is improved.
The rotary rocket part takes propellant as working medium, and the propellant is bipropellant, is not limited to bipropellant, and can be monopropellant or tripropellant, etc. The bipropellant can be ignited by spontaneous combustion after collision, an additional ignition system is not needed, the weight of the system can be reduced, and the working reliability of the system can be improved.
Further, as shown in fig. 10, in order to ensure that the fuel oil delivered to the rotary thrust chamber is uniformly distributed in the circumferential direction, axial oil grooves 510 uniformly distributed in the circumferential direction are formed on the inner surfaces of the middle layer thin-wall section and the outer layer thin-wall section of the main shaft 501.
It will be understood by those skilled in the art that the foregoing is only a preferred embodiment of the present invention, and is not intended to limit the invention, and although the invention has been described in detail with reference to the foregoing examples, it will be apparent to those skilled in the art that various changes in the form and details of the embodiments may be made and equivalents may be substituted for elements thereof. All modifications, equivalents and the like which come within the spirit and principle of the invention are intended to be included within the scope of the invention.
Claims (5)
1. A turbine rotary rocket combined engine, which is characterized by comprising a turbine engine part and a rotary rocket part which are connected along the axial direction;
the turbine engine part comprises an air inlet volute, an air compressor, a rotor shaft, a flame tube and a gas turbine;
the rotary rocket part comprises a lateral fuel oil sealing cavity, a lateral oxidant sealing cavity, a front bearing seat, a bearing sleeve, a combustion chamber body, a combustion chamber jet pipe, a afterburning chamber casing and a thrust jet pipe;
the afterburning chamber casing is fixedly connected to the air inlet volute; the thrust spray pipe is connected with the afterburning chamber casing; the combustion generator consisting of a lateral fuel oil sealing cavity, a lateral oxidant sealing cavity, a front bearing seat, a bearing sleeve, a combustion chamber body and a combustion chamber spray pipe is positioned in a afterburning chamber formed by the afterburning chamber casing;
the combustion chamber body comprises a main shaft extending out and a combustion chamber cavity arranged in the main shaft; the main shaft is a stepped concentric shaft sleeve shaft and comprises an inner layer solid section, a middle layer thin-wall section and an outer layer thin-wall section which are sleeved together; a first fuel flow channel is formed in the middle-layer thin-wall section, a first oxidant flow channel is formed in the outer-layer thin-wall section, and the first fuel flow channel and the first oxidant flow channel are both annular; the second fuel flow channels are uniformly distributed along the radial direction of the tail end of the first fuel flow channel, and the first fuel flow channels and the second fuel flow channels form fuel flow channels; the oxidant flow channels II are uniformly distributed along the radial direction of the tail end of the oxidant flow channel I, and the oxidant flow channel I and the oxidant flow channel II form an oxidant flow channel; the combustion chamber cavity is positioned between the plane on which the second fuel flow channel is distributed and the plane on which the second oxidant flow channel is distributed, one end of the second fuel flow channel is communicated with the first fuel flow channel, and the other end of the second fuel flow channel is communicated with the combustion chamber cavity through a fuel nozzle; one end of the oxidant flow channel II is communicated with the oxidant flow channel I, and the other end of the oxidant flow channel II is communicated with the combustion chamber cavity through an oxidant nozzle; the rear end face of the combustion chamber cavity is uniformly provided with a plurality of inclined exhaust outlets, and a combustion chamber spray pipe is arranged in each exhaust outlet;
the lateral fuel oil sealing cavity, the lateral oxidant sealing cavity, the front bearing seat and the bearing sleeve are sequentially and fixedly connected along the axis direction of the main shaft and sleeved on the main shaft; the lateral fuel oil seal cavity is positioned between the inner-layer solid section and the middle-layer thin-wall section of the main shaft, and is provided with a fuel oil inlet along the radial direction;
and after the oxygen-rich high-temperature fuel gas generated by the turbine engine part and the oil-rich high-temperature fuel gas generated by the fuel gas generator of the rotary rocket part are combusted again in the afterburning chamber, the fuel gas is sprayed out through the thrust spray pipe to generate thrust.
2. The turbine rotary rocket combination engine according to claim 1, wherein the inner surfaces of the middle layer thin-wall section and the outer layer thin-wall section of the main shaft are both provided with axial oil grooves which are uniformly distributed along the circumferential direction.
3. A turbo-rocket motor assembly according to claim 1, wherein the cross-sectional diameter of the internal cavity of the combustor nozzle is first reduced and then increased.
4. A turbo-rocket motor assembly according to claim 1, wherein said number of nozzles is 3.
5. A turbo-rotary rocket assembly engine according to claim 1 wherein each of said second fuel flow passage and said second oxidizer flow passage is six.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210759257.XA CN115199438B (en) | 2022-06-29 | 2022-06-29 | Turbine rotary rocket combined engine |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210759257.XA CN115199438B (en) | 2022-06-29 | 2022-06-29 | Turbine rotary rocket combined engine |
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| Publication Number | Publication Date |
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| CN115199438A true CN115199438A (en) | 2022-10-18 |
| CN115199438B CN115199438B (en) | 2024-09-13 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN202210759257.XA Active CN115199438B (en) | 2022-06-29 | 2022-06-29 | Turbine rotary rocket combined engine |
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE4227907A1 (en) * | 1992-08-22 | 1994-02-24 | Kurt Mannl | Liquid-fuelled rocket injection system - has rotary casing with ports slid pneumatically or hydraulically so as to vary free cross-section of injection ports |
| JP2005147122A (en) * | 2003-10-09 | 2005-06-09 | Hiroyasu Tanigawa | Various rocket combined engine |
| US20080256925A1 (en) * | 2007-04-17 | 2008-10-23 | Pratt & Whitney Rocketdyne, Inc. | Compact, high performance swirl combustion rocket engine |
| CN107503862A (en) * | 2017-10-10 | 2017-12-22 | 北京航空航天大学 | A kind of hybrid rocket combination circulation propulsion system and its control method |
| CN112253335A (en) * | 2020-10-16 | 2021-01-22 | 中国科学院力学研究所 | Gas generator for driving turbine in rocket engine |
-
2022
- 2022-06-29 CN CN202210759257.XA patent/CN115199438B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE4227907A1 (en) * | 1992-08-22 | 1994-02-24 | Kurt Mannl | Liquid-fuelled rocket injection system - has rotary casing with ports slid pneumatically or hydraulically so as to vary free cross-section of injection ports |
| JP2005147122A (en) * | 2003-10-09 | 2005-06-09 | Hiroyasu Tanigawa | Various rocket combined engine |
| US20080256925A1 (en) * | 2007-04-17 | 2008-10-23 | Pratt & Whitney Rocketdyne, Inc. | Compact, high performance swirl combustion rocket engine |
| CN107503862A (en) * | 2017-10-10 | 2017-12-22 | 北京航空航天大学 | A kind of hybrid rocket combination circulation propulsion system and its control method |
| CN112253335A (en) * | 2020-10-16 | 2021-01-22 | 中国科学院力学研究所 | Gas generator for driving turbine in rocket engine |
Non-Patent Citations (1)
| Title |
|---|
| 潘宏亮;林彬彬;刘洋;: "加力式空气涡轮火箭发动机特性研究", 固体火箭技术, vol. 33, no. 06, 15 December 2010 (2010-12-15), pages 650 - 655 * |
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| Publication number | Publication date |
|---|---|
| CN115199438B (en) | 2024-09-13 |
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