CN115450798A - Long-time working solid rocket engine and long tail jet pipe thereof - Google Patents
Long-time working solid rocket engine and long tail jet pipe thereof Download PDFInfo
- Publication number
- CN115450798A CN115450798A CN202211345759.4A CN202211345759A CN115450798A CN 115450798 A CN115450798 A CN 115450798A CN 202211345759 A CN202211345759 A CN 202211345759A CN 115450798 A CN115450798 A CN 115450798A
- Authority
- CN
- China
- Prior art keywords
- carbon ring
- section
- long
- combustion chamber
- hole
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
- 239000007787 solid Substances 0.000 title claims abstract description 52
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical group [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 claims abstract description 186
- 238000009413 insulation Methods 0.000 claims abstract description 41
- 238000002485 combustion reaction Methods 0.000 claims description 59
- 238000002347 injection Methods 0.000 claims description 37
- 239000007924 injection Substances 0.000 claims description 37
- 230000002829 reductive effect Effects 0.000 claims description 26
- VYPSYNLAJGMNEJ-UHFFFAOYSA-N Silicium dioxide Chemical compound O=[Si]=O VYPSYNLAJGMNEJ-UHFFFAOYSA-N 0.000 claims description 14
- 238000009434 installation Methods 0.000 claims description 14
- 239000002131 composite material Substances 0.000 claims description 12
- 230000000149 penetrating effect Effects 0.000 claims description 11
- 239000000463 material Substances 0.000 claims description 10
- 229920000049 Carbon (fiber) Polymers 0.000 claims description 7
- 239000004917 carbon fiber Substances 0.000 claims description 7
- 230000007774 longterm Effects 0.000 claims description 7
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims description 7
- 239000000377 silicon dioxide Substances 0.000 claims description 7
- CREMABGTGYGIQB-UHFFFAOYSA-N carbon carbon Chemical compound C.C CREMABGTGYGIQB-UHFFFAOYSA-N 0.000 claims description 6
- 239000011203 carbon fibre reinforced carbon Substances 0.000 claims description 6
- 229910052799 carbon Inorganic materials 0.000 claims description 5
- SBYXRAKIOMOBFF-UHFFFAOYSA-N copper tungsten Chemical compound [Cu].[W] SBYXRAKIOMOBFF-UHFFFAOYSA-N 0.000 claims description 5
- 230000007423 decrease Effects 0.000 claims description 4
- 230000008595 infiltration Effects 0.000 claims description 4
- 238000001764 infiltration Methods 0.000 claims description 4
- ISWSIDIOOBJBQZ-UHFFFAOYSA-N phenol group Chemical group C1(=CC=CC=C1)O ISWSIDIOOBJBQZ-UHFFFAOYSA-N 0.000 claims description 4
- 239000007789 gas Substances 0.000 description 37
- 229920006335 epoxy glue Polymers 0.000 description 8
- 239000002737 fuel gas Substances 0.000 description 8
- -1 phenolic aldehyde Chemical class 0.000 description 6
- 230000000903 blocking effect Effects 0.000 description 4
- 239000007921 spray Substances 0.000 description 4
- 238000002679 ablation Methods 0.000 description 3
- RYGMFSIKBFXOCR-UHFFFAOYSA-N Copper Chemical compound [Cu] RYGMFSIKBFXOCR-UHFFFAOYSA-N 0.000 description 2
- 229910052802 copper Inorganic materials 0.000 description 2
- 239000010949 copper Substances 0.000 description 2
- 239000004744 fabric Substances 0.000 description 2
- 230000000670 limiting effect Effects 0.000 description 2
- 238000009991 scouring Methods 0.000 description 2
- 238000004804 winding Methods 0.000 description 2
- 239000002759 woven fabric Substances 0.000 description 2
- 238000010521 absorption reaction Methods 0.000 description 1
- 238000004364 calculation method Methods 0.000 description 1
- 238000006243 chemical reaction Methods 0.000 description 1
- 230000003628 erosive effect Effects 0.000 description 1
- 238000011010 flushing procedure Methods 0.000 description 1
- 238000012423 maintenance Methods 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 230000036961 partial effect Effects 0.000 description 1
- 230000002035 prolonged effect Effects 0.000 description 1
- 230000001681 protective effect Effects 0.000 description 1
- 239000003870 refractory metal Substances 0.000 description 1
- 238000007789 sealing Methods 0.000 description 1
- 238000004088 simulation Methods 0.000 description 1
- 239000000243 solution Substances 0.000 description 1
- WFKWXMTUELFFGS-UHFFFAOYSA-N tungsten Chemical compound [W] WFKWXMTUELFFGS-UHFFFAOYSA-N 0.000 description 1
- 229910052721 tungsten Inorganic materials 0.000 description 1
- 239000010937 tungsten Substances 0.000 description 1
- 238000005406 washing 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
- F02K9/97—Rocket nozzles
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/97—Rocket nozzles
- F02K9/974—Nozzle- linings; Ablative coatings
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Cylinder Crankcases Of Internal Combustion Engines (AREA)
Abstract
The application relates to solid rocket engine technical field especially relates to a solid rocket engine of long-time work and long tail nozzle thereof, includes: the long tail nozzle comprises a long tail nozzle shell, a convergent equal straight section heat insulation layer, a back lining, a front carbon ring, a throat liner, a rear carbon ring and a plug cover. The throat area of long tail nozzle can be guaranteed for a long time by the aid of the long tail nozzle, the thermal protection capacity is greatly improved, the working reliability is guaranteed, and the working time of the long tail nozzle can reach more than 200 s.
Description
Technical Field
The application relates to the technical field of solid rocket engines, in particular to a solid rocket engine working for a long time and a long tail nozzle thereof.
Background
The solid rocket engine has the advantages of simple structure, convenient maintenance, high reliability and simple and convenient operation, and is widely applied to missile power systems. The jet pipe is an energy conversion device of a solid rocket engine, and can convert the heat energy of high-temperature and high-pressure fuel gas into kinetic energy so as to generate thrust. The long tail nozzle is a long straight pipe with a section of equal section added at a proper position in the middle of a conventional Laval nozzle, and can be used on a tactical missile which adopts a solid rocket engine and has high maneuverability requirement. On one hand, the installation space of the missile control system component can be reserved by adopting the long straight pipe with the equal section; on the other hand, the center of mass of the solid rocket engine can be moved forwards, and meanwhile, the change of the center of mass position of the missile adopting the solid rocket engine in the flight process is reduced as much as possible, so that the control of the missile is facilitated.
The main functions of the nozzle/long tail nozzle are as follows: a) Controlling the flow of the high-temperature and high-pressure gas through the throat area, keeping the high-temperature and high-pressure gas in a combustion chamber of the solid rocket engine at a preset pressure, and keeping normal combustion of the charge in the combustion chamber; b) The heat energy of the high-temperature and high-pressure gas is converted into kinetic energy, so that the thrust required by the flying of the missile is realized; c) Bearing the high temperature and high pressure of the fuel gas during the operation of the solid rocket engine. Therefore, the spray pipe/long tail spray pipe works in high-temperature and high-pressure environment for a long time, and the working environment is very severe.
However, since the long nozzle operates in a high-temperature and high-pressure environment for a long time, the temperature of the wall surface of the long nozzle needs to be strictly controlled to ensure that the missile control system components mounted around the long straight pipe can normally operate. Meanwhile, the length of the equal section is increased by the long-tail nozzle, and the flowing time of high-temperature and high-pressure gas in the long-tail nozzle is prolonged, so that the difficulty of maintaining the throat area and heat protection of the long-tail nozzle is higher compared with that of a conventional nozzle. In order to ensure that the throat area of the long-tail nozzle is basically kept unchanged under the condition of long-time high-temperature and high-pressure gas flushing, the throat lining positioned in the nozzle is usually made of refractory metals (such as tungsten copper infiltration), but the heat transfer coefficient of the materials is very large, so that the requirement on the thermal protection of the long-tail nozzle is more severe, and a proper working environment can be provided for missile control system components. At present, the working time of the existing long tail nozzle is usually about 100s, which is difficult to meet the requirement of cruise missile which needs longer time (more than 200 s).
In view of this, in order to further broaden the applications of the solid rocket engine, it is a technical problem that those skilled in the art need to solve so far that the working time of the solid rocket engine is more than 200 s.
Disclosure of Invention
The application provides a solid rocket engine of long time work and long tail nozzle thereof, can guarantee the throat area of long tail nozzle for a long time, has still promoted hot protective capacities by a wide margin, has ensured operational reliability, makes its operating time can reach more than 200 s.
In order to solve the technical problem, the application provides the following technical scheme:
a long nozzle for a long-life solid rocket engine, comprising: the long tail nozzle comprises a long tail nozzle shell, a convergent equal straight section heat insulation layer, a back lining, a front carbon ring, a throat liner and a rear carbon ring; the rear carbon ring is provided with a rear carbon ring mounting air injection hole which penetrates through two ends of the rear carbon ring; the throat liner is provided with throat liner air injection holes penetrating through two ends, the throat liner is inserted into and fixed to the front section of the rear carbon ring mounting air injection hole, the throat liner is made of tungsten copper infiltration, and the thermal conductivity of the rear carbon ring is lower than that of the throat liner; the front carbon ring is provided with a front carbon ring jet hole penetrating through two ends, the rear section of the front carbon ring is inserted into and fixed to the front section of the throat liner jet hole, the outer surface of the front section of the front carbon ring is in contact with the outer surface of the front section of the rear carbon ring, so that the throat liner is wrapped by the front carbon ring and the rear carbon ring, and the thermal conductivity of the front carbon ring is lower than that of the throat liner; the backing has a backing mounting hole through both ends, the front carbon ring and the rear carbon ring carrying throat insert are inserted and fixed into the backing mounting hole, and the rear carbon ring heat insulation groove on the outer surface of the rear carbon ring is made to contain air, the backing is inserted and fixed to the rear section of the long tail nozzle housing, and the backing heat insulation groove on the outer surface of the backing is made to contain air; the convergent isopipe heat insulating layer is provided with a convergent isopipe heat insulating layer air injection hole penetrating through two ends, the convergent isopipe heat insulating layer is inserted into and fixed to the front section of the long tail nozzle shell, and the rear end face of the convergent isopipe heat insulating layer is in contact with and fixed to the front end face of the back lining and the front end face of the front carbon ring.
The long tail nozzle of the solid rocket engine working for a long time as described above, wherein it is preferable that the diameter of the front section of the rear carbon ring mounting gas injection hole is reduced from front to rear, and the diameter of the rear section of the rear carbon ring mounting gas injection hole is increased from front to rear; the outer diameter of the throat insert is reduced from front to back so as to be arranged at the front section of the rear carbon ring installation air injection hole.
The long tail nozzle of the solid rocket engine which operates for a long time as described above, wherein, preferably, the diameter of the part of the front section of the rear carbon ring mounting gas injection hole, which is close to the combustion chamber, is constant; the diameter of the middle part of the front section of the rear carbon ring mounting gas injection hole is gradually reduced from front to back; the diameter of the part of the front section of the rear carbon ring mounting gas injection hole, which is far away from the combustion chamber, is unchanged; the diameter of the rear section of the rear carbon ring mounting gas injection hole is gradually increased from front to back; the outer diameter of the part, close to the combustion chamber, of the front section of the throat insert is unchanged so as to be in contact with the part, close to the combustion chamber, of the front section of the rear carbon ring mounting gas injection hole; the outer diameter of the part, far away from the combustion chamber, of the front section of the throat insert is gradually reduced from front to back so as to be in contact with the middle part of the front section of the rear carbon ring mounting fumarole; the outer diameter of the rear section of the throat insert is unchanged so as to be in contact with the part, far away from the combustion chamber, of the front section of the rear carbon ring installation gas injection hole.
The long tail nozzle of the solid rocket engine working for a long time as described above, wherein preferably, the diameter of the front section of the throat insert jet hole is reduced from front to back, and the diameter of the rear section of the throat insert jet hole is increased from front to back; the outer diameter of the rear section of the front carbon ring is gradually reduced from front to back so as to be arranged at the front section of the throat liner fumarole.
The long tail nozzle of a solid rocket engine operating for a long time as described above, wherein it is preferable that the diameter of the rear section of the backing installation hole is gradually reduced from front to rear, and the outer diameter of the rear section of the rear carbon ring is gradually reduced from front to rear to contact with the rear section of the backing installation hole; the front section of the back lining mounting hole is a second-order stepped hole, and the diameter of the first-order stepped hole far away from the combustion chamber is smaller than that of the first-order stepped hole close to the combustion chamber; the outer diameter of the front section of the front carbon ring is unchanged, the outer diameter of the front section of the rear carbon ring is unchanged, and the outer surfaces of the front section of the front carbon ring and the front section of the rear carbon ring are in contact with a first-order step hole of the backing mounting hole, which is far away from the combustion chamber; the end surface of the rear section of the heat insulating layer of the convergent equal-straight section is contacted with a step hole of the backing mounting hole close to the combustion chamber.
The long tail pipe of the solid rocket motor working for a long time as described above, wherein, preferably, the rear section of the heat insulating layer of the convergent isopipe is a three-step stage, and the diameter of the one-step stage far away from the combustion chamber is smaller than that of the one-step stage in the middle, and the diameter of the one-step stage in the middle is smaller than that of the one-step stage near the combustion chamber; the stage far away from the combustion chamber is contacted with the part, close to the combustion chamber, of the front section of the front carbon ring gas injection hole of the front carbon ring, the stage in the middle is contacted with the stage one hole, close to the combustion chamber, of the back lining, and the stage one step section close to the combustion chamber is contacted with the inner surface of the front section of the long tail nozzle shell.
The long-tail nozzle of the solid rocket engine working for a long time as described above, wherein, preferably, the rear carbon ring is made of a piercing carbon-carbon composite material, the front carbon ring is made of a piercing carbon-carbon composite material, the backing is made of a high silica/phenolic material, and the convergent equal straight section heat insulation layer is made of a 2.5D carbon fiber integrally woven composite material.
The long tail nozzle of the solid rocket engine working for a long time as described above, wherein preferably, a plurality of rear carbon ring heat insulation grooves are uniformly distributed in the circumferential direction on the outer surface of the front section of the rear carbon ring; two groups of backing heat insulation grooves which are uniformly distributed in the circumferential direction are arranged on the outer surface of the backing; the first group of back lining heat insulation grooves correspond to the front section of the rear carbon ring, and the second group of back lining heat insulation grooves correspond to the front section of the front carbon ring.
The long nozzle of the solid rocket engine operating for a long time as described above, wherein the blanking cap is preferably adhesively fixed to the rear end face of the rear carbon ring by the nozzle opening of the divergent section of the long nozzle casing.
A long-term solid rocket engine, wherein the nozzle is the long tail nozzle of the long-term solid rocket engine.
Has the advantages that:
1. the equal straight section heat insulation layer of convergence adopts the whole combined material that weaves of 2.5D carbon fiber in this application, and this material is wholly woven 2.5D inline profile modeling woven fabric/phenolic aldehyde as the ablation resistant layer with the 12K carbon fiber to high silica cloth/phenolic aldehyde are as insulating layer complex winding, can adapt to the washing away of high temperature high pressure gas under the long-time work, compares the carbon fiber/phenolic aldehyde moulded products that the tradition adopted in a large number, has promoted the operational reliability of spray tube.
2. According to past engineering experience and simulation calculation, the position of the throat insert and the front end of the throat insert are the position with the highest heat exchange coefficient in the spray pipe, and the thermal protection at the position is particularly important for a solid rocket engine working for a long time, so that the front end of the throat insert is provided with a front carbon ring for puncturing a carbon/carbon composite material to effectively reduce the heat exchange amount at the position, and the rear end of the throat insert is provided with a rear carbon ring for puncturing the carbon/carbon composite material to adapt to the thermal load brought by the large heat conductivity of tungsten copper infiltration, and simultaneously, the throat insert can bear the scouring of high-temperature and high-pressure fuel gas as an expansion heat insulation layer.
3. In this application back carbocycle and back lining fitting surface, back lining and long tail nozzle casing fitting surface are provided with a plurality of heat-insulating slots, under the prerequisite of guaranteeing structural strength, the air thermal conductivity in the heat-insulating slot is very low (only about 5% of high silica/phenolic aldehyde), thereby evenly set up the heat-insulating slot can greatly reduced local area's heat conduction in the throat lining footpath, ensure that here long tail nozzle casing is unlikely to the too high structure that leads to of temperature and became invalid and for arranging in long tail nozzle guided missile control system part all around provides suitable operational environment, especially to the long tail nozzle of operating time more than 200s, can further improve the operational reliability of long tail nozzle.
Drawings
In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the drawings needed to be used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments described in the present application, and other drawings can be obtained by those skilled in the art according to the drawings.
FIG. 1 is a schematic illustration of a long tail nozzle of a long-term operating solid rocket engine provided in an embodiment of the present application;
FIG. 2 is an enlarged view of portion A of FIG. 1;
110-long tail pipe shell, 120-convergent section heat insulation layer, 130-backing, 131-backing heat insulation groove, 140-front carbon ring, 150-throat liner, 160-rear carbon ring, 161-rear carbon ring heat insulation groove and 170-plug cover.
Detailed Description
Reference will now be made in detail to embodiments of the present application, examples of which are illustrated in the accompanying drawings, wherein like or similar reference numerals refer to the same or similar elements or elements having the same or similar function throughout. The embodiments described below with reference to the accompanying drawings are exemplary only for explaining the present application and are not construed as limiting the present application.
In addition, spatial relationship terms such as "upper", "lower", "front", "rear", and the like are used for convenience of description to explain a positional relationship between two members. In the present application, "front" refers to a direction toward the combustion chamber, and "rear" refers to a direction away from the combustion chamber.
Referring to fig. 1 and 2, the present application provides a long nozzle for a long-term operation of a solid rocket engine, comprising: a long jet casing 110, a convergent isopipe insulation 120, a backing 130, a forward carbon ring 140, a throat insert 150, and an aft carbon ring 160.
The rear carbon ring 160 has a rear carbon ring mounting air hole penetrating through both ends, and a front section of the rear carbon ring mounting air hole is used for fixing the throat insert 150. Optionally, the diameter of the front section of the rear carbon ring mounting gas orifice is reduced from front to back, and the diameter of the rear section of the rear carbon ring mounting gas orifice is increased from front to back. Optionally, the diameter of the front section of the rear carbon ring mounting gas injection hole close to the combustion chamber is unchanged, and the rear carbon ring mounting gas injection hole is used for being in contact with the outer surface of the front section of the throat insert 150 close to the combustion chamber; the diameter of the middle part of the front section of the rear carbon ring mounting gas injection hole is gradually reduced from front to back and is used for being in contact with the outer surface of the part, far away from the combustion chamber, of the front section of the throat insert 150; the diameter of the part, far away from the combustion chamber, of the front section of the rear carbon ring installation gas injection hole is unchanged, and the rear carbon ring installation gas injection hole is used for being in contact with the outer surface of the rear section of the throat insert 150; the diameter of the rear section of the rear carbon ring mounting gas injection hole is gradually increased from front to back, and the rear carbon ring mounting gas injection hole is used for injecting high-temperature and high-pressure fuel gas. Optionally, the outer diameter of the front section of the rear carbon ring 160 is constant, and the outer diameter of the rear section of the rear carbon ring 160 is gradually reduced from front to rear.
Optionally, the thermal conductivity of the rear carbon ring 160 is lower than the thermal conductivity of the throat insert 150, so as to effectively reduce the amount of heat exchange of the throat insert 150 with the rear carbon ring 160. Optionally, the rear carbon ring 160 is made of a piercing carbon-carbon composite material to adapt to the thermal load caused by the large thermal conductivity of the tungsten-infiltrated copper, and meanwhile, the rear carbon ring 160 serving as an expansion section heat insulation layer can also bear the erosion of high-temperature and high-pressure fuel gas.
The throat insert 150 has a throat insert gas orifice penetrating both ends, and the throat insert 150 is inserted into and fixed to the front section of the rear carbon ring installation gas orifice. Optionally, the throat insert 150 is adhesively fixed to the front section of the rear carbon ring mounting fumarole by epoxy glue. Optionally, the outer diameter of the throat insert 150 decreases from front to back to fit in the front section of the rear carbon ring mounting jet hole. Optionally, the outer diameter of the part of the front section of the throat insert 150 close to the combustion chamber is unchanged, and the throat insert is used for contacting with the part of the front section of the rear carbon ring installation fumarole close to the combustion chamber; the outer diameter of the part, far away from the combustion chamber, of the front section of the throat insert 150 is gradually reduced from front to back and is used for being in contact with the middle part of the front section of the rear carbon ring mounting gas injection hole; the outer diameter of the rear section of the throat insert 150 is constant for contact with the portion of the front section of the aft carbon ring mounting fumarole away from the combustion chamber. Optionally, the diameter of the front section of the throat insert gas orifice is reduced from front to back, and the diameter of the rear section of the throat insert gas orifice is increased from front to back. Still optionally, the throat insert gas orifice surface is an arcuate surface.
Optionally, the throat insert 150 is made of a tungsten copper infiltrated material to ensure that the throat area does not change within the working time of more than 200 seconds, thereby ensuring that the performance of the solid rocket engine does not decrease.
The front carbon ring 140 has a front carbon ring air injection hole penetrating both ends, and the rear section of the front carbon ring 140 is inserted into and fixed to the front section of the throat insert air injection hole, and the outer surface of the front section of the front carbon ring 140 is brought into contact with the outer surface of the front section of the rear carbon ring 160, thereby wrapping the throat insert 150 by the front carbon ring 140 and the rear carbon ring 160. Optionally, the rear section of the front carbon ring 140 is adhesively fixed to the front section of the throat insert gas orifice by epoxy glue. Optionally, the outer diameter of the front section of the front carbon ring 140 is unchanged, and the outer diameter of the rear section of the front carbon ring 140 is gradually reduced from front to rear so as to be mounted on the front section of the throat insert gas orifice. Optionally, the outer surface of the front section of the front carbon ring 140 is connected to the outer surface of the rear section of the front carbon ring 140 through a radial connecting surface, and is in contact with the front end surface of the rear carbon ring 160 through the radial connecting surface, so that the throat insert 150 is wrapped by the front carbon ring 140 and the rear carbon ring 160, and direct contact between the throat insert 150 and the backing 130 is avoided. Optionally, the rear section of the front carbon ring 140 is inserted into a partial position of the front section of the throat insert gas orifice to avoid the front carbon ring 140 completely covering the throat of the throat insert 150.
Optionally, the diameter of the front carbon ring gas injection holes decreases from front to back. Still alternatively, the portion of the forward carbon ring injector hole forward adjacent the combustion chamber is tapered in diameter to form an inclined surface to contact the portion of the rear face of the converging straight section insulation layer 120.
Optionally, the thermal conductivity of the front carbon ring 140 is lower than the thermal conductivity of the throat insert 150 to effectively reduce the amount of heat exchange at the front section local position of the throat insert 150. Optionally, the front carbon ring 140 is made of a piercing carbon-carbon composite material, and the thermal conductivity of the front carbon ring is only about one fourth of that of tungsten infiltrated copper, so that the heat exchange amount of the front section local position of the throat insert 150 can be effectively reduced, and the working reliability of the long-tail nozzle is improved.
The backing 130 has a backing mounting hole through both ends, and the front carbon ring 140 and the rear carbon ring 160 are inserted and fixed into the backing mounting hole carrying the throat insert 150. Optionally, the front carbon ring 140 and the rear carbon ring 160 are adhesively fixed into the backing mounting hole by epoxy glue. Alternatively, the diameter of the rear section of the backing mounting hole is gradually reduced from front to rear to contact the outer surface of the rear section of the rear carbon ring 160. Further alternatively, the front section of the backing mounting hole is a second step hole, and the diameter of the first step hole far from the combustion chamber is smaller than that of the first step hole near the combustion chamber, so that the first step hole far from the combustion chamber is in contact with the outer surfaces of the front section of the rear carbon ring 160 and the front section of the front carbon ring 140, and the first step hole near the combustion chamber is in contact with the convergent isobaric insulating layer 120.
Optionally, the backing 130 is made of a high silica/phenolic material, that is, the backing 130 is a high silica/phenolic molded article, which has the characteristics of low density and thermal conductivity, excellent heat absorption capability, ablation resistance, and thermal stability. Also optionally, a concave rear carbon ring heat insulation groove 161 is provided on the outer surface of the rear carbon ring 160. Preferably, a concave rear carbon ring heat insulation groove 161 is formed on the outer surface of the front section of the rear carbon ring 160 to correspond to the throat insert 150, so that after the rear carbon ring 160 is installed in the backing installation hole and the outer surface of the rear carbon ring 160 contacts with the inner surface of the backing installation hole, a certain amount of air is enclosed in the rear carbon ring heat insulation groove 161, and the thermal conductivity of the air in the rear carbon ring heat insulation groove 161 is far lower than that of the rear carbon ring 160, thereby effectively reducing the heat exchange amount in the region during the operation of the long tail nozzle. Optionally, a plurality of rear carbon ring heat insulation grooves 161 are uniformly distributed in the circumferential direction on the outer surface of the front section of the rear carbon ring 160.
The backing 130 is inserted into and fixed to the rear section of the long jet nozzle housing 110. Optionally, the backing 130 is adhesively secured within the jet nozzle housing 110 by an epoxy glue. Optionally, a concave backing insulation slot 131 is provided on the outer surface of the backing 130, such that after the backing 130 is installed in the long jet nozzle housing 110 and the outer surface of the backing 130 is in contact with the inner surface of the long jet nozzle housing 110, a certain amount of air is enclosed in the backing insulation slot 131, and the thermal conductivity of the air in the backing insulation slot 131 is much lower than the thermal conductivity of the backing 130, thereby effectively reducing the amount of heat exchange in this area during the operation of the long jet nozzle. Still alternatively, two sets of backing insulation slots 131 are provided on the outer surface of the backing 130, evenly distributed circumferentially. Optionally, the first set of backing insulation slots corresponds to the front section of the rear carbon ring 160, and the second set of backing insulation slots corresponds to the front section of the front carbon ring 140.
The convergent isopipe insulating layer 120 has a convergent isopipe insulating layer gas orifice penetrating both ends, the convergent isopipe insulating layer 120 is inserted and fixed to the front section of the long jet nozzle housing 110, and the rear end face of the convergent isopipe insulating layer 120 is fixed in contact with the front end face of the backing 130 and the front end face of the front carbon ring 140. Preferably, the convergent isopipe insulation layer 120 is adhesively secured into the long jet nozzle housing 110 by epoxy glue, and the rear end face of the convergent isopipe insulation layer 120 is adhesively secured to the front end face of the backing 130 and the front end face of the front carbon ring 140 by epoxy glue. Optionally, the rear section of the convergent-isocandela insulation layer 120 is a three-step stage, and the diameter of the stage far from the combustion chamber is smaller than that of the stage in the middle, and the diameter of the stage in the middle is smaller than that of the stage near the combustion chamber, so that the stage far from the combustion chamber contacts with the part of the front section of the front carbon ring gas injection hole of the front carbon ring 140 near the combustion chamber, the stage in the middle contacts with the stage hole of the backing 130 near the combustion chamber, and the stage near the combustion chamber contacts with the inner surface of the front section of the long tail nozzle housing 110. Optionally, the diameter of the stepped section of the convergent isocandela insulation layer 120 away from the combustion chamber is gradually reduced from front to back, thereby forming an inclined surface to contact the inclined surface of the combustion chamber near the front section of the front carbon ring gas orifice of the front carbon ring 140.
Optionally, the material of the heat insulating layer 120 at the convergent equal straight section is a 2.5D carbon fiber integrally-woven composite material, the material uses 12K carbon fiber integrally-woven 2.5D in-line profiling woven fabric/phenolic aldehyde as an ablation resistant layer, and uses high silica cloth/phenolic aldehyde as a heat insulating layer for composite winding, so that the material has very strong scouring resistance and heat insulating property.
On the basis of the above, the blanking cap 170 is adhesively fixed to the rear end surface of the rear carbon ring 160 from the nozzle of the expanded section of the long jet nozzle housing 110 to close the solid rocket engine, thereby ensuring that the interior of the solid rocket engine is in a dry environment. Optionally, the front end face of the blocking cover 170 is adhesively fixed to the outer end face of the rear carbon ring 160 by epoxy glue.
The application also provides a long-time working solid rocket engine which comprises the long tail nozzle of the long-time working solid rocket engine.
When the long tail nozzle of the solid rocket engine working for a long time is assembled, the throat liner 150 is firstly installed and fixed into the rear carbon ring 160, the front carbon ring 140 is installed and fixed into the throat liner 150, then the whole body formed by the front carbon ring 140, the throat liner 150 and the rear carbon ring 160 is installed and fixed into the backing 130, then the whole body formed by the throat liner 150, the rear carbon ring 160, the front carbon ring 140 and the backing 130 is installed and fixed into the long tail nozzle shell 110, then the convergent equal-straight-section heat insulation layer 120 is installed and fixed into the long tail nozzle shell 110, and finally the blocking cover 170 is installed and fixed onto the rear end face of the rear carbon ring 160, so that the sealing reliability of the long tail nozzle is ensured, and the assembly of the long tail nozzle of the solid rocket engine working for a long time is completed. Wherein, each part is bonded by epoxy glue and is pressed and cured by a tool.
When the solid rocket engine working for a long time is in a working state, high-temperature and high-pressure fuel gas generated in the combustion chamber flows through the convergent equal-section heat insulating layer 120, the front carbon ring 140, the throat liner 150 and the rear carbon ring 160, the high-temperature and high-pressure fuel gas forces the blocking cover 170 to be opened, the blocking cover 170 and the high-temperature and high-pressure fuel gas are flushed out immediately, the solid rocket engine working for a long time starts to work normally until the charge burnout work is finished, and a long tail nozzle of the solid rocket engine working for a long time can keep complete and effective thermal protection within more than 200s of normal work, so that a proper working environment can be provided for missile control system components arranged around the long tail nozzle, and the application of the solid rocket engine is widened.
It will be evident to those skilled in the art that the present application is not limited to the details of the foregoing illustrative embodiments, and that the present application may be embodied in other specific forms without departing from the spirit or essential attributes thereof. The present embodiments are therefore to be considered in all respects as illustrative and not restrictive, the scope of the application being indicated by the appended claims rather than by the foregoing description, and all changes which come within the meaning and range of equivalency of the claims are therefore intended to be embraced therein. Any reference sign in a claim should not be construed as limiting the claim concerned.
Furthermore, it should be understood that although the present specification describes embodiments, not every embodiment includes only a single embodiment, and such description is for clarity purposes only, and it is to be understood that all embodiments may be combined as appropriate by one of ordinary skill in the art to form other embodiments as will be apparent to those of skill in the art from the description herein.
Claims (10)
1. A long tail nozzle for a long-term operation solid rocket engine, comprising: the long tail nozzle comprises a long tail nozzle shell, a convergent equal straight section heat insulation layer, a back lining, a front carbon ring, a throat liner and a rear carbon ring;
the rear carbon ring is provided with a rear carbon ring mounting air injection hole which penetrates through two ends of the rear carbon ring;
the throat liner is provided with throat liner air injection holes penetrating through two ends, the throat liner is inserted into and fixed to the front section of the rear carbon ring mounting air injection hole, the throat liner is made of tungsten copper infiltration, and the thermal conductivity of the rear carbon ring is lower than that of the throat liner;
the front carbon ring is provided with a front carbon ring jet hole penetrating through two ends, the rear section of the front carbon ring is inserted into and fixed to the front section of the throat liner jet hole, the outer surface of the front section of the front carbon ring is in contact with the outer surface of the front section of the rear carbon ring, so that the throat liner is wrapped by the front carbon ring and the rear carbon ring, and the thermal conductivity of the front carbon ring is lower than that of the throat liner;
the back lining is provided with a back lining mounting hole penetrating through two ends, the front carbon ring and the rear carbon ring carrying throat linings are inserted and fixed into the back lining mounting hole, the rear carbon ring heat insulation groove on the outer surface of the rear carbon ring is used for containing air, the back lining is inserted and fixed to the rear section of the long tail nozzle shell, and the back lining heat insulation groove on the outer surface of the back lining is used for containing air;
the convergent isopipe heat insulating layer is provided with a convergent isopipe heat insulating layer air injection hole penetrating through two ends, the convergent isopipe heat insulating layer is inserted into and fixed to the front section of the long tail nozzle shell, and the rear end face of the convergent isopipe heat insulating layer is in contact with and fixed to the front end face of the back lining and the front end face of the front carbon ring.
2. The long tail nozzle of a long-life solid rocket engine according to claim 1, wherein the diameter of the front section of the rear carbon ring mounting jet hole decreases from front to rear, and the diameter of the rear section of the rear carbon ring mounting jet hole increases from front to rear;
the outer diameter of the throat insert is reduced from front to back so as to be installed at the front section of the rear carbon ring installation fumarole.
3. The long tail nozzle of a long-life solid rocket engine according to claim 1, wherein the diameter of the front section of the rear carbon ring-mounted jet hole near the combustion chamber is constant; the diameter of the middle part of the front section of the rear carbon ring mounting gas injection hole is gradually reduced from front to back; the diameter of the part of the front section of the rear carbon ring mounting gas injection hole, which is far away from the combustion chamber, is unchanged; the diameter of the rear section of the rear carbon ring mounting gas injection hole is gradually increased from front to back;
the outer diameter of the part, close to the combustion chamber, of the front section of the throat insert is unchanged so as to be in contact with the part, close to the combustion chamber, of the front section of the rear carbon ring mounting gas injection hole; the outer diameter of the part, far away from the combustion chamber, of the front section of the throat insert is gradually reduced from front to back so as to be in contact with the middle part of the front section of the rear carbon ring mounting gas injection hole; the outer diameter of the rear section of the throat insert is unchanged so as to be in contact with the part, far away from the combustion chamber, of the front section of the rear carbon ring mounting gas injection hole.
4. A long tail nozzle of a long-term solid rocket engine according to any one of claims 1 to 3, wherein the diameter of the front section of the throat insert jet hole is reduced from front to back, and the diameter of the rear section of the throat insert jet hole is increased from front to back;
the outer diameter of the rear section of the front carbon ring is gradually reduced from front to back so as to be arranged at the front section of the throat insert gas orifice.
5. The longnozzle of a solid rocket engine working for a long period of time according to any one of claims 1 to 3, wherein the diameter of the rear section of the backing installation hole is gradually reduced from front to rear, and the outer diameter of the rear section of the rear carbon ring is gradually reduced from front to rear to contact with the rear section of the backing installation hole;
the front section of the back lining mounting hole is a second-order stepped hole, and the diameter of the first-order stepped hole far away from the combustion chamber is smaller than that of the first-order stepped hole close to the combustion chamber;
the outer diameter of the front section of the front carbon ring is unchanged, the outer diameter of the front section of the rear carbon ring is unchanged, and the outer surfaces of the front section of the front carbon ring and the front section of the rear carbon ring are in contact with a first-order step hole of the backing mounting hole, which is far away from the combustion chamber;
the end surface of the rear section of the heat insulating layer of the convergent equal straight section is contacted with a step hole of the backing mounting hole, which is close to the combustion chamber.
6. The long tail pipe of a long-term solid rocket motor according to claim 5 wherein the rear section of the convergent-iso-straight section insulation layer is a three-step stage, and the diameter of the one-step stage away from the combustion chamber is smaller than the diameter of the one-step stage in the middle, and the diameter of the one-step stage in the middle is smaller than the diameter of the one-step stage near the combustion chamber;
the stage far away from the combustion chamber is contacted with the part, close to the combustion chamber, of the front section of the front carbon ring gas injection hole of the front carbon ring, the stage in the middle is contacted with the stage one hole, close to the combustion chamber, of the back lining, and the stage one step section close to the combustion chamber is contacted with the inner surface of the front section of the long tail nozzle shell.
7. The long tail nozzle of a solid rocket engine working for a long time according to any one of claims 1 to 3, wherein the rear carbon ring is made of a piercing carbon-carbon composite material, the front carbon ring is made of a piercing carbon-carbon composite material, the backing is made of a high silica/phenolic material, and the convergent straight section heat insulating layer is made of a 2.5D carbon fiber integrally woven composite material.
8. The long tail nozzle of a solid rocket engine working for a long time according to any one of claims 1 to 3, wherein a plurality of rear carbon ring heat insulation grooves are uniformly distributed in the circumferential direction on the outer surface of the front section of the rear carbon ring;
two groups of backing heat insulation grooves which are uniformly distributed in the circumferential direction are arranged on the outer surface of the backing; the first group of back lining heat insulation grooves correspond to the front section of the rear carbon ring, and the second group of back lining heat insulation grooves correspond to the front section of the front carbon ring.
9. The long nozzle of a solid rocket engine working for an extended period of time as claimed in any one of claims 1 to 3, wherein the blanking cap is adhesively secured to the rear end face of the rear carbon ring by the nozzle opening of the divergent section of the long nozzle housing.
10. A long-life solid rocket engine, characterized in that the nozzle is a long tail nozzle of the long-life solid rocket engine according to any one of claims 1 to 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211345759.4A CN115450798A (en) | 2022-10-31 | 2022-10-31 | Long-time working solid rocket engine and long tail jet pipe thereof |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211345759.4A CN115450798A (en) | 2022-10-31 | 2022-10-31 | Long-time working solid rocket engine and long tail jet pipe thereof |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN115450798A true CN115450798A (en) | 2022-12-09 |
Family
ID=84311661
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211345759.4A Pending CN115450798A (en) | 2022-10-31 | 2022-10-31 | Long-time working solid rocket engine and long tail jet pipe thereof |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115450798A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117028079A (en) * | 2023-10-08 | 2023-11-10 | 北京星河动力装备科技有限公司 | Jet device, rocket power device and rocket |
Citations (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2849860A (en) * | 1955-10-17 | 1958-09-02 | Norton Co | Rocket motor with recrystallized silicon carbide throat insert |
| US4643356A (en) * | 1977-08-31 | 1987-02-17 | United Technologies Corporation | Cooling liner for convergent-divergent exhaust nozzle |
| FR2811018A1 (en) * | 2000-06-30 | 2002-01-04 | Alliant Techsystems Inc | Rocket motor assembly insulation or thermal protection ablation material is made from impregnated resin matrix with carbonizing reinforcement |
| US6711901B1 (en) * | 2000-01-21 | 2004-03-30 | Alliant Techsystems Inc. | Rocket motor nozzle assemblies having vacuum plasma-sprayed refractory metal shell throat inserts, methods of making, and rocket motors including same |
| CN202596924U (en) * | 2012-04-18 | 2012-12-12 | 湖北航天技术研究院总体设计所 | Sectioned long exhaust nozzle structure for solid rocket engine |
| CN204493010U (en) * | 2014-12-09 | 2015-07-22 | 上海新力动力设备研究所 | A kind of Combination backing structure of solid propellant rocket jet pipe larynx lining |
| CN105736184A (en) * | 2014-12-09 | 2016-07-06 | 上海新力动力设备研究所 | Large thrust ratio, long-working micro-ablation throat insert and throat structure of expansion section |
| CN106050477A (en) * | 2016-07-28 | 2016-10-26 | 湖北航天技术研究院总体设计所 | Combined throat liner spraying pipe of solid rocket engine and manufacturing method |
| CN106979095A (en) * | 2017-05-08 | 2017-07-25 | 湖北航天技术研究院总体设计所 | The integrated jet pipe and its manufacture method of a kind of global formation |
| CN107091169A (en) * | 2017-06-29 | 2017-08-25 | 湖北三江航天江河化工科技有限公司 | A kind of rocket engine |
| CN109139301A (en) * | 2018-09-07 | 2019-01-04 | 西安航天化学动力厂 | A kind of integrated solid rocket motor nozzle of thermal protection structure |
| CN109736975A (en) * | 2018-11-30 | 2019-05-10 | 西安航天动力技术研究所 | A kind of jet pipe and its design method reducing the outer wall temperature of expansion segment |
| CN110805505A (en) * | 2019-11-15 | 2020-02-18 | 西安近代化学研究所 | Heat insulation layer assembly suitable for aluminum alloy long-tail nozzle solid rocket engine |
| CN111075606A (en) * | 2019-12-27 | 2020-04-28 | 北京中科宇航探索技术有限公司 | Solid rocket engine charging structure with adjustable combustion surface ratio and rocket engine |
| CN114198223A (en) * | 2021-11-29 | 2022-03-18 | 湖北航泰科技有限公司 | One-step curing molding full-composite engine spray pipe |
-
2022
- 2022-10-31 CN CN202211345759.4A patent/CN115450798A/en active Pending
Patent Citations (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2849860A (en) * | 1955-10-17 | 1958-09-02 | Norton Co | Rocket motor with recrystallized silicon carbide throat insert |
| US4643356A (en) * | 1977-08-31 | 1987-02-17 | United Technologies Corporation | Cooling liner for convergent-divergent exhaust nozzle |
| US6711901B1 (en) * | 2000-01-21 | 2004-03-30 | Alliant Techsystems Inc. | Rocket motor nozzle assemblies having vacuum plasma-sprayed refractory metal shell throat inserts, methods of making, and rocket motors including same |
| FR2811018A1 (en) * | 2000-06-30 | 2002-01-04 | Alliant Techsystems Inc | Rocket motor assembly insulation or thermal protection ablation material is made from impregnated resin matrix with carbonizing reinforcement |
| CN202596924U (en) * | 2012-04-18 | 2012-12-12 | 湖北航天技术研究院总体设计所 | Sectioned long exhaust nozzle structure for solid rocket engine |
| CN105736184A (en) * | 2014-12-09 | 2016-07-06 | 上海新力动力设备研究所 | Large thrust ratio, long-working micro-ablation throat insert and throat structure of expansion section |
| CN204493010U (en) * | 2014-12-09 | 2015-07-22 | 上海新力动力设备研究所 | A kind of Combination backing structure of solid propellant rocket jet pipe larynx lining |
| CN106050477A (en) * | 2016-07-28 | 2016-10-26 | 湖北航天技术研究院总体设计所 | Combined throat liner spraying pipe of solid rocket engine and manufacturing method |
| CN106979095A (en) * | 2017-05-08 | 2017-07-25 | 湖北航天技术研究院总体设计所 | The integrated jet pipe and its manufacture method of a kind of global formation |
| CN107091169A (en) * | 2017-06-29 | 2017-08-25 | 湖北三江航天江河化工科技有限公司 | A kind of rocket engine |
| CN109139301A (en) * | 2018-09-07 | 2019-01-04 | 西安航天化学动力厂 | A kind of integrated solid rocket motor nozzle of thermal protection structure |
| CN109736975A (en) * | 2018-11-30 | 2019-05-10 | 西安航天动力技术研究所 | A kind of jet pipe and its design method reducing the outer wall temperature of expansion segment |
| CN110805505A (en) * | 2019-11-15 | 2020-02-18 | 西安近代化学研究所 | Heat insulation layer assembly suitable for aluminum alloy long-tail nozzle solid rocket engine |
| CN111075606A (en) * | 2019-12-27 | 2020-04-28 | 北京中科宇航探索技术有限公司 | Solid rocket engine charging structure with adjustable combustion surface ratio and rocket engine |
| CN114198223A (en) * | 2021-11-29 | 2022-03-18 | 湖北航泰科技有限公司 | One-step curing molding full-composite engine spray pipe |
Non-Patent Citations (3)
| Title |
|---|
| 熊文波;刘宇;任军学;于泉;司学龙;: "小推力长时间工作固体火箭发动机初步试验研究", 固体火箭技术, no. 04, 15 August 2009 (2009-08-15), pages 413 - 417 * |
| 王相宇等: "固体发动机壳体瞬态温度场的数值模拟研究", 中国航天第三专业信息网第三十九届技术交流会暨第三届空天动力联合会议论文集——S05发动机热管理技术, 22 August 2018 (2018-08-22), pages 35 - 42 * |
| 胡春波;张钢锤;何国强;姜立东;彭永林;: "受激光控制分解式固体微推力器内流动特性研究", 固体火箭技术, no. 02, 25 June 2006 (2006-06-25), pages 17 - 20 * |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117028079A (en) * | 2023-10-08 | 2023-11-10 | 北京星河动力装备科技有限公司 | Jet device, rocket power device and rocket |
| CN117028079B (en) * | 2023-10-08 | 2024-02-20 | 北京星河动力装备科技有限公司 | Jet device, rocket power device and rocket |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN111810318B (en) | Single-chamber double-thrust solid rocket engine and rocket | |
| US6151887A (en) | Combustion chamber for rocket engine | |
| CN115450798A (en) | Long-time working solid rocket engine and long tail jet pipe thereof | |
| CN202596924U (en) | Sectioned long exhaust nozzle structure for solid rocket engine | |
| CN110080909B (en) | Jet pipe of solid rocket engine | |
| CN109139301A (en) | A kind of integrated solid rocket motor nozzle of thermal protection structure | |
| CN210714882U (en) | Rocket engine jet pipe | |
| CN117028079B (en) | Jet device, rocket power device and rocket | |
| CN112761823B (en) | Expansion section lateral drainage thrust vector control spray pipe | |
| EP3239476B1 (en) | Case clearance control system and corresponding gas turbine engine | |
| US3712063A (en) | Cooled pintle assembly | |
| RU2403491C2 (en) | Thermal power cooled wall construction of high-temperature air-gas path element | |
| CN109707539B (en) | Integrated composite spray pipe based on gradient material | |
| CN111237087A (en) | Novel active and passive micropore plate composite cooling structure for aerospace power and cooling method | |
| CN115573829A (en) | Ablation-resistant structure of expansion section outlet | |
| CN108087153A (en) | A kind of Ducted rocket with cooling combination | |
| CN218407625U (en) | Solid rocket engine jet pipe with integrated throat liner and expansion section | |
| US4681261A (en) | Heat resistant short nozzle | |
| CN109736975B (en) | Spray pipe capable of reducing temperature of outer wall of expansion section and design method thereof | |
| CN208106596U (en) | A kind of solid rocket ramjet jet pipe | |
| CN111396218A (en) | Engine jet pipe, solid rocket engine and solid rocket | |
| CN114876673B (en) | Low-cost ablation-resistant embedded spray pipe and processing method thereof | |
| RU2171388C2 (en) | Chamber of liquid-prpellant rocket engine | |
| CN218467729U (en) | Long-time low-thrust spray pipe diffusion section | |
| CN109515757A (en) | A kind of jet aircraft attitude-control device |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |