CN111963339B - Liquid film cooling rail attitude control engine thrust chamber - Google Patents
Liquid film cooling rail attitude control engine thrust chamber Download PDFInfo
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- CN111963339B CN111963339B CN202010837309.1A CN202010837309A CN111963339B CN 111963339 B CN111963339 B CN 111963339B CN 202010837309 A CN202010837309 A CN 202010837309A CN 111963339 B CN111963339 B CN 111963339B
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- 239000007788 liquid Substances 0.000 title claims abstract description 142
- 238000001816 cooling Methods 0.000 title claims abstract description 20
- 238000002347 injection Methods 0.000 claims abstract description 177
- 239000007924 injection Substances 0.000 claims abstract description 177
- 239000003380 propellant Substances 0.000 claims abstract description 106
- 238000002485 combustion reaction Methods 0.000 claims abstract description 105
- 239000007921 spray Substances 0.000 claims abstract description 26
- 238000004088 simulation Methods 0.000 claims description 9
- 238000004364 calculation method Methods 0.000 claims description 6
- 230000007480 spreading Effects 0.000 claims description 6
- 238000004458 analytical method Methods 0.000 claims description 3
- 238000012938 design process Methods 0.000 claims description 3
- 238000006243 chemical reaction Methods 0.000 abstract description 2
- 239000000243 solution Substances 0.000 description 5
- 230000008901 benefit Effects 0.000 description 3
- 238000000034 method Methods 0.000 description 3
- 230000000694 effects Effects 0.000 description 2
- 230000006872 improvement Effects 0.000 description 2
- 230000008569 process Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 239000000110 cooling liquid Substances 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000002737 fuel gas Substances 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
- 230000001172 regenerating effect Effects 0.000 description 1
- 238000005507 spraying Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
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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
- F02K9/62—Combustion or thrust chambers
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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
- F02K9/44—Feeding propellants
- F02K9/52—Injectors
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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
- F02K9/44—Feeding propellants
- F02K9/56—Control
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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
- F02K9/96—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by specially adapted arrangements for testing or measuring
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- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
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- General Engineering & Computer Science (AREA)
- Combustion Methods Of Internal-Combustion Engines (AREA)
Abstract
The invention discloses a liquid film cooling rail attitude control engine thrust chamber, wherein a first propellant chamber is arranged on the outer side of the head part of the thrust chamber, a second propellant chamber is arranged on the outer side of the body part of the thrust chamber and positioned in the middle position; the wall of the head part of the thrust chamber is provided with a plurality of first spray holes communicated with the first propellant chamber; a plurality of combustion injection holes and a plurality of liquid film injection holes are uniformly formed in the wall of the thrust chamber body part along the circumferential direction; the combustion jetting holes and the liquid film jetting holes respectively jet second propellants to the head part and the throat part of the thrust chamber, and the liquid film jetting is enabled to face the throat part of the thrust chamber, and a tangential liquid inlet mode is adopted, so that the liquid film loss can be reduced, and the cooling performance of the liquid film can be more fully exerted; meanwhile, the combustion injection holes face the head of the thrust chamber, a tangential liquid inlet mode is adopted, the cooling performance of the liquid film can be exerted, and meanwhile, the reaction with the first propellant can also be realized, so that the proportion of the liquid film in the propellant can be reduced, and the whole body cooling of the thrust chamber is realized.
Description
Technical Field
The invention relates to an engine thrust chamber, in particular to a liquid film cooling rail attitude control engine thrust chamber.
Background
The improvement of the specific impact performance of the orbit attitude control engine can effectively prolong the service life of the spacecraft or increase the quality of effective load, and can bring remarkable economic benefit and military benefit. The main factors restricting the improvement of the specific impulse performance of the rail attitude control engine are as follows: high performance injector technology, high temperature resistant material processes and inefficient utilization of liquid films. Among them, inefficient use of the liquid film has become the primary limiting factor in the reduction of the specific thrust of the engine.
In a conventional orbital attitude control engine, there are two ways of liquid film cooling:
firstly, the propellant is sprayed to the inner wall surface of the thrust chamber through a small hole with a certain inclination angle arranged at the edge of the spraying disc, thereby forming protection to the wall surface.
And secondly, a centrifugal injector is adopted, a conical liquid film is formed after the propellant leaves the injector and is sputtered to the inner wall surface of the thrust chamber to form protection.
However, both of the above methods have the following problems:
(1) when the propellant used as the liquid film is sprayed to the inner wall surface, the liquid drops are rebounded due to the impact angle between the propellant and the inner wall surface, so that the actual amount of the cooling liquid film is reduced;
(2) the thermal environment of the area near the throat of the thrust chamber is the worst, which is the important of the thermal protection design, but in the prior scheme, the liquid film enters from the head of the thrust chamber, and because the liquid film is far away from the throat, the liquid film amount is lost due to the unstable flow of the surface of the liquid film and the entrainment effect of fuel gas in the downstream flowing process of the wall surface.
Both of the above reasons cause the liquid film ratio to be high. In order to improve the specific impulse performance of the engine, the key of the problem is how to reduce the proportion of the liquid film as much as possible while ensuring the cooling.
Disclosure of Invention
In order to solve the problem that the specific impact performance of an engine is reduced due to the fact that the proportion of a liquid film in a thrust chamber of an existing rail attitude control engine is high in the background art, the invention provides a liquid film cooling rail attitude control engine thrust chamber.
The specific technical scheme of the invention is as follows:
the invention provides a liquid film cooling rail attitude control engine thrust chamber, wherein a first propellant chamber is arranged on the outer side of the head part of the thrust chamber, a second propellant chamber is arranged on the outer side of the body part of the thrust chamber and positioned in the middle position; the wall of the head part of the thrust chamber is provided with a plurality of first spray holes communicated with the first propellant chamber; a plurality of combustion injection holes and a plurality of liquid film injection holes are uniformly formed in the wall of the thrust chamber body part along the circumferential direction;
the plurality of combustion injection holes and the plurality of liquid film injection holes are communicated with the second propellant cavity, and the combustion injection holes and the liquid film injection holes are alternately arranged;
the combustion injection hole is tangential to the wall of the body part of the thrust chamber, inclines towards the head part of the thrust chamber, and is sprayed out through the combustion injection hole to adhere to the wall and swirl and moves towards the head part of the thrust chamber; the propellant sprayed out from all the combustion injection holes is consistent with the rotational flow direction of the wall of the thrust chamber body part;
the liquid film injection hole is tangent to the wall of the body part of the thrust chamber, inclines towards the throat part of the thrust chamber, and moves towards the throat part of the thrust chamber while the propellant sprayed out of the liquid film injection hole is attached to the wall of the rotational flow; the propellant sprayed out from all the liquid film spray holes is consistent along the rotational flow direction of the wall of the body part of the thrust chamber.
Furthermore, in order to avoid backflow, annular bulges are arranged on the inner wall of the thrust chamber between the outlets of the combustion injection holes and the outlets of the liquid film injection holes.
Further, in order to further improve the utilization rate of the propellant, the swirling direction of the propellant sprayed from all the combustion injection holes along the wall of the thrust chamber body part is consistent with the swirling direction of the propellant sprayed from all the liquid film injection holes along the wall of the thrust chamber body part.
Further, the thrust chamber has an ellipse-like structure as a whole in cross section.
Furthermore, an included angle between the combustion injection hole and a radial section of the head part of the thrust chamber is alpha, alpha is more than or equal to 6 degrees and is more than or equal to 2 degrees, and an included angle between the liquid film injection hole and a radial section of the throat part of the thrust chamber is beta; beta is more than or equal to 1 degree at 3 degrees.
Further, the number of the combustion injection holes is in direct proportion to the size of the included angle alpha; the number of the liquid film injection holes is in direct proportion to the size of the included angle beta.
Further, the ratio of the flow rate of the second propellant sprayed through all the combustion injection holes to the flow rate of the second propellant sprayed through all the liquid film injection holes is 85%: 15% to 95%: 5 percent.
Furthermore, the injection pressure drop of the second propellant chamber is 20-40% of the pressure of the thrust chamber.
Further, the specific design process of the liquid film cooling rail attitude control engine thrust chamber is as follows:
step 1: respectively calculating the flow of the second propellant sprayed by all the combustion jetting holes and the flow of the propellant sprayed by all the liquid film jetting holes according to the total flow of the required second propellant, the flow of the second propellant sprayed by all the combustion jetting holes and the flow ratio of the second propellant sprayed by all the liquid film jetting holes;
step 2: respectively calculating the flow areas of all combustion injection holes and all liquid film injection holes according to the result of the step 1 and the injection pressure drop of the second propellant cavity, wherein the specific calculation formula is as follows:
wherein A is the flow area of all combustion injection holes and all liquid film injection holes;
the flow rate of the propellant sprayed from all the combustion spray holes or the flow rate of the propellant sprayed from all the liquid film spray holes; cdThe flow coefficient is set to be 0.7-0.8, rho is the density of the second propellant, and delta p is the jetting pressure drop of the initially set second propellant cavity;
and step 3: preliminarily establishing a thrust chamber model in simulation software;
step 3.1: according to the known overall dimension of the thrust chamber, an initially set angle alpha and an initially set angle beta, preliminarily determining the number of combustion injection holes according to the principle that one combustion injection hole is arranged every 10-20 mm in the circumferential direction of the inner wall of the thrust chamber; determining the number of liquid film injection holes according to the principle that the number of the combustion injection holes is integral multiple of the number of the liquid film injection holes, and requiring uniform parts of the combustion injection holes and the liquid film injection holes along the circumferential direction and alternately arranging the combustion injection holes and the liquid film injection holes;
step 3.2: respectively calculating the aperture of each combustion injection hole and the aperture of each liquid film injection hole according to the flow areas of all the combustion injection holes and all the liquid film injection holes calculated in the step 2 and the number of the combustion injection holes and the liquid film injection holes preliminarily determined in the step 3.1, so as to preliminarily establish a thrust chamber model;
the specific calculation formula of the aperture of the combustion injection hole and the aperture of the liquid film injection hole is as follows:
d is the aperture of the combustion injection holes or the liquid film injection holes, and n is the number of the combustion injection holes or the liquid film injection holes;
and 4, step 4: analyzing by simulation software to obtain a final thrust chamber model;
step 4.1: simulation software carries out simulation analysis on the initially established thrust chamber model, whether the injection pressure drop of the second propellant cavity obtained through simulation is matched with the injection pressure drop of the second propellant cavity initially set in the step 2 is compared, if not, the step 2 is skipped, and the flow coefficient C is modifieddValue taking, repeating the rest steps until the injection pressure drop obtained by simulation is matched with the injection pressure drop initially set, and skipping to the step 4.2;
step 4.2: and (3) checking the spreading uniformity of the liquid film on the inner wall of the thrust chamber, if the uniformity is poor, skipping to the step 3, adjusting the number of the combustion injection holes and the number of the liquid film injection holes, and repeating the rest steps until the good spreading uniformity of the liquid film is obtained, so that the final thrust chamber model is determined.
Compared with the existing thrust chamber liquid film cooling scheme, the thrust chamber liquid film cooling method has the beneficial effects that:
according to the invention, the body part of the thrust chamber is additionally provided with the second propellant chamber, and the second propellant is respectively sprayed to the head part and the tail part of the thrust chamber through the combustion injection hole and the liquid film injection hole; meanwhile, the combustion injection holes face the head of the thrust chamber, a tangential liquid inlet mode is adopted, the cooling performance of the liquid film can be exerted, and the reaction with the first propellant can be realized, so that the proportion of the liquid film in the propellant can be reduced, the whole-body cooling of the thrust chamber is perfectly realized, and the specific impact performance of the engine and the utilization rate of the propellant are effectively improved.
Drawings
Fig. 1 is a structural view of a thrust chamber.
Fig. 2 is a schematic view of the angle α.
Fig. 3 is a schematic view of the included angle β.
The reference numbers are as follows:
1-thrust chamber head, 2-first propellant chamber, 3-thrust chamber body, 4-second propellant chamber, 5-first spray orifice, 6-combustion spray orifice, 7-liquid film spray orifice, 8-thrust chamber throat and 9-annular bulge.
Detailed Description
The technical solutions of the present invention will be described clearly and completely with reference to the accompanying drawings, and it should be understood that the described embodiments are some, but not all embodiments of the present invention. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, are within the scope of the present invention.
In the description of the present invention, it should be noted that the terms "first" and "second" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
In the description of the present invention, it should also be noted that, unless otherwise explicitly specified or limited, the terms "mounted," "connected," and "connected" are to be construed broadly, e.g., as meaning either a fixed connection, a removable connection, or an integral connection; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meaning of the above terms in the present invention can be understood as appropriate by those of ordinary skill in the art.
The embodiment provides a specific structure of a liquid film cooling rail attitude control engine thrust chamber, as shown in fig. 1, a first propellant chamber 2 is arranged on the outer side of a head part 1 of the thrust chamber, a body part 3 of the thrust chamber is arranged on the outer side of a middle position, and a second propellant chamber 4 is arranged on the outer side of the middle position; the wall of the thrust chamber head 1 is provided with a plurality of first spray holes 5 communicated with the first propellant chamber 2; a plurality of combustion injection holes 6 and a plurality of liquid film injection holes 7 are uniformly arranged on the wall of the thrust chamber body part 3 along the circumferential direction;
the plurality of combustion injection holes 6 and the plurality of liquid film injection holes 7 are communicated with the second propellant chamber 4, and the combustion injection holes 6 and the liquid film injection holes 7 are alternately arranged;
as shown in fig. 2, the combustion injection hole is tangential to the wall of the body part of the thrust chamber, and the combustion injection hole is inclined towards the head part of the thrust chamber (i.e. an included angle α exists between the combustion injection hole and a radial section of the head part of the thrust chamber, for convenience of understanding, the radial section of the head part of the thrust chamber is equivalent to a straight line a in fig. 2), and the propellant sprayed through the combustion injection hole adheres to the wall and swirls and moves towards the head part of the thrust chamber; the propellant sprayed out from all the combustion injection holes is consistent with the rotational flow direction of the wall of the thrust chamber body part;
as shown in fig. 3, the liquid film injection hole is tangential to the wall of the body part of the thrust chamber, and the liquid film injection hole is inclined toward the throat part of the thrust chamber (i.e., an included angle exists between the combustion injection hole and a radial section of the throat part of the thrust chamber, β is for easy understanding, the radial sections of the throat part of the thrust chamber are all equivalent to a straight line a in fig. 3), and the propellant sprayed from the liquid film injection hole moves toward the throat part of the thrust chamber while swirling against the wall; the propellant sprayed out from all the liquid film spray holes is consistent along the rotational flow direction of the wall of the body part of the thrust chamber.
For ease of understanding, fig. 2 and 3 each equate the radial cross-section of the thrust chamber head and the thrust chamber throat to a straight line a.
When in work: the first propellant is injected into the first propellant cavity 2 and is sprayed into the thrust chamber through the first spray holes 5, the second propellant is injected into the second propellant cavity 4, and the second propellant has certain spray pressure drop because the second propellant cavity 4 has certain spray pressure drop, so the second propellant is sprayed into the thrust chamber through the combustion spray holes 6 and the liquid film spray holes 7, wherein the second propellant sprayed from the combustion spray holes 6 enters the combustion chamber and then is attached to the inner wall in a counter-flow upward manner, the second propellant has good heat protection effect on the front section of the body part of the combustion chamber and can be regarded as a liquid film, and then the second propellant from the counter-flow upward manner is atomized and combusted with the first propellant, which is similar to the 'regenerative cooling' of a large-thrust engine. The second propellant sprayed by the liquid film injection holes 7 enters the combustion chamber and then flows down along the inner wall to protect the rear section and the throat part of the combustion chamber with larger heat load, so that the structure realizes the whole body cooling of the combustion chamber, effectively reduces the liquid film loss, reduces the proportion of the liquid film which does not participate in the combustion, and improves the specific impulse performance of the combustion chamber.
In the embodiment, an included angle between the combustion injection hole and a radial section of the head part of the thrust chamber is alpha, the included angle is more than or equal to 6 degrees and more than or equal to 2 degrees, and an included angle between the liquid film injection hole and the inclination of the radial section of the throat part of the thrust chamber is beta; beta is more than or equal to 1 degree at 3 degrees, and the number of the combustion injection holes 6 is in direct proportion to the size of the included angle alpha; the number of the liquid film injection holes 7 is in direct proportion to the size of the included angle beta (namely, when the number of the injection holes is more, the angle can be a high value, and when the number of the injection holes is less, the angle is a low value). The proportion of the flow of the second propellant sprayed from all the combustion spray holes 6 to the flow of the second propellant sprayed from all the liquid film spray holes 7 is selected to be 85%: 15% to 95%: 5 percent; the injection pressure drop of the second propellant chamber 4 is about 20-40% of the pressure of the thrust chamber. The number of the combustion injection holes is 6-12, and is integral multiple of the number of the liquid film injection holes, and the number of the liquid film injection holes is 3-6 in the embodiment.
The thrust chamber structure in the present embodiment also provides the following optimized design:
1. because a plurality of combustion injection holes 6 and a plurality of liquid film injection holes 7 are attached to the wall and swirl, in order to avoid backflow, annular bulges 9 with proper height are arranged on the inner wall of the thrust chamber between the outlets of the combustion injection holes 6 and the liquid film injection holes 7.
2. In order to further improve the utilization rate of the propellant, the swirling direction of the propellant sprayed from all the combustion injection holes 6 along the inner wall of the thrust chamber body 3 is consistent with the swirling direction of the propellant sprayed from all the liquid film injection holes 7 along the inner wall of the thrust chamber body 3.
3. The section of the thrust chamber is of an ellipse-like structure as a whole. The structure has the advantage that the second propellant entering the body part 3 of the thrust chamber through the combustion injection hole 6 can easily reach the head part 1 of the thrust chamber and be mixed and combusted with the first propellant.
The number, the aperture and the inclination angle of the combustion injection holes 6 and the liquid film injection holes 7 are different necessarily due to different thrust chambers, so the embodiment also gives a specific design process of the thrust chamber:
step 1: respectively calculating the flow of the second propellant sprayed by all the combustion jetting holes and the flow of the propellant sprayed by all the liquid film jetting holes according to the total flow of the required second propellant, the flow of the second propellant sprayed by all the combustion jetting holes and the flow ratio of the second propellant sprayed by all the liquid film jetting holes;
step 2: respectively calculating the flow areas of all combustion injection holes and all liquid film injection holes according to the result of the step 1 and the injection pressure drop of the second propellant cavity, wherein the specific calculation formula is as follows:
wherein A is the flow area of all combustion injection holes and all liquid film injection holes;
the flow rate of the propellant sprayed from all the combustion spray holes or the flow rate of the propellant sprayed from all the liquid film spray holes; cdThe flow coefficient is set to be 0.7-0.8, rho is the density of the second propellant, and delta p is the jetting pressure drop of the initially set second propellant cavity;
and step 3: preliminarily establishing a thrust chamber model in simulation software;
step 3.1: according to the known overall dimension of the thrust chamber, an initially set angle alpha and an initially set angle beta, preliminarily determining the number of combustion injection holes according to the principle that one combustion injection hole is arranged every 10-20 mm in the circumferential direction of the inner wall of the thrust chamber; determining the number of liquid film injection holes according to the principle that the number of the combustion injection holes is integral multiple of the number of the liquid film injection holes, and requiring uniform parts of the combustion injection holes and the liquid film injection holes along the circumferential direction and alternately arranging the combustion injection holes and the liquid film injection holes;
step 3.2: respectively calculating the aperture of each combustion injection hole and the aperture of each liquid film injection hole according to the flow areas of all the combustion injection holes and all the liquid film injection holes calculated in the step 2 and the number of the combustion injection holes and the liquid film injection holes preliminarily determined in the step 3.1, so as to preliminarily establish a thrust chamber model;
the specific calculation formula of the aperture of the combustion injection hole and the aperture of the liquid film injection hole is as follows:
d is the aperture of the combustion injection holes or the liquid film injection holes, and n is the number of the combustion injection holes or the liquid film injection holes;
and 4, step 4: analyzing by simulation software to obtain a final thrust chamber model;
step 4.1: simulation software carries out simulation analysis on the initially established thrust chamber model, whether the injection pressure drop of the second propellant cavity obtained through simulation is matched with the injection pressure drop of the second propellant cavity initially set in the step 2 is compared, if not, the step 2 is skipped, and the flow coefficient C is modifieddValue taking, repeating the rest steps until the injection pressure drop obtained by simulation is matched with the injection pressure drop initially set, and skipping to the step 4.2;
step 4.2: and (3) checking the spreading uniformity of the liquid film on the inner wall of the thrust chamber, if the uniformity is poor, skipping to the step 3, adjusting the number of the combustion injection holes and the number of the liquid film injection holes, and repeating the rest steps until the good spreading uniformity of the liquid film is obtained, so that the final thrust chamber model is determined.
Finally, it should be noted that: the above embodiments are only used to illustrate the technical solution of the present invention, and not to limit the same; while the invention has been described in detail and with reference to the foregoing embodiments, it will be understood by those skilled in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some or all of the technical features may be equivalently replaced; and the modifications or the substitutions do not make the essence of the corresponding technical solutions depart from the scope of the technical solutions of the embodiments of the present invention.
Claims (9)
1. The utility model provides a liquid film cooling rail appearance accuse engine thrust room which characterized in that:
a first propellant cavity is arranged on the outer side of the head part of the thrust chamber, a second propellant cavity is arranged on the outer side of the body part of the thrust chamber and positioned in the middle position;
the wall of the head part of the thrust chamber is provided with a plurality of first spray holes communicated with the first propellant chamber; a plurality of combustion injection holes and a plurality of liquid film injection holes are uniformly formed in the wall of the thrust chamber body part along the circumferential direction;
the plurality of combustion injection holes and the plurality of liquid film injection holes are communicated with the second propellant cavity and are alternately arranged;
the combustion injection hole is tangential to the wall of the body part of the thrust chamber, inclines towards the head part of the thrust chamber, and is sprayed out through the combustion injection hole to adhere to the wall and swirl and moves towards the head part of the thrust chamber; the propellant sprayed out from all the combustion injection holes is consistent with the rotational flow direction of the wall of the thrust chamber body part;
the liquid film injection hole is tangent to the wall of the body part of the thrust chamber, inclines towards the throat part of the thrust chamber, and moves towards the throat part of the thrust chamber while the propellant sprayed out of the liquid film injection hole is attached to the wall of the rotational flow; the propellant sprayed out through all the liquid film spray holes is consistent along the rotational flow direction of the wall of the thrust chamber body part;
and annular bulges are arranged on the inner wall of the thrust chamber between the outlets of the combustion injection holes and the outlets of the liquid film injection holes.
2. The liquid film cooled rail attitude control engine thrust chamber of claim 1, characterized in that: the swirling direction of the propellant sprayed from all the combustion injection holes along the wall of the thrust chamber body part is consistent with the swirling direction of the propellant sprayed from all the liquid film injection holes along the wall of the thrust chamber body part.
3. The liquid film cooled rail attitude control engine thrust chamber of claim 2, characterized in that: the section of the thrust chamber is of an ellipse-like structure as a whole.
4. The liquid film cooled rail attitude control engine thrust chamber of claim 3, characterized in that: the included angle between the combustion injection hole and the radial section of the head part of the thrust chamber is alpha, the included angle is more than or equal to alpha and more than or equal to 2 degrees at 6 degrees, and the included angle between the liquid film injection hole and the radial section of the throat part of the thrust chamber is beta; beta is more than or equal to 1 degree at 3 degrees.
5. The liquid film cooled rail attitude controlled engine thrust chamber of claim 4, wherein: the number of the combustion injection holes is in direct proportion to the size of the included angle alpha; the number of the liquid film injection holes is in direct proportion to the size of the included angle beta.
6. The liquid film cooled rail attitude controlled engine thrust chamber of claim 5, wherein: the proportion of the flow of the second propellant sprayed from all the combustion spray holes to the flow of the second propellant sprayed from all the liquid film spray holes is 85%: 15% to 95%: 5 percent.
7. The liquid film cooled rail attitude controlled engine thrust chamber of claim 6, wherein: the injection pressure drop of the second propellant cavity is 20-40% of the pressure of the thrust chamber.
8. The liquid film cooled rail attitude controlled engine thrust chamber of claim 7, wherein: the specific design process is as follows:
step 1: respectively calculating the flow of the second propellant sprayed by all the combustion jetting holes and the flow of the propellant sprayed by all the liquid film jetting holes according to the total flow of the required second propellant, the flow of the second propellant sprayed by all the combustion jetting holes and the flow ratio of the second propellant sprayed by all the liquid film jetting holes;
step 2: respectively calculating the flow areas of all combustion injection holes and all liquid film injection holes according to the result of the step 1 and the injection pressure drop of the second propellant cavity, wherein the specific calculation formula is as follows:
wherein A is the flow area of all combustion injection holes and all liquid film injection holes;
the flow rate of the propellant sprayed from all the combustion spray holes or the flow rate of the propellant sprayed from all the liquid film spray holes; cdThe flow coefficient is set to be 0.7-0.8, rho is the density of the second propellant, and delta p is the jetting pressure drop of the initially set second propellant cavity;
and step 3: preliminarily establishing a thrust chamber model in simulation software;
step 3.1: according to the known overall dimension of the thrust chamber, an initially set angle alpha and an initially set angle beta, preliminarily determining the number of combustion injection holes according to the principle that one combustion injection hole is arranged every 10-20 mm in the circumferential direction of the inner wall of the thrust chamber; determining the number of liquid film injection holes according to the principle that the number of the combustion injection holes is integral multiple of the number of the liquid film injection holes, and requiring uniform parts of the combustion injection holes and the liquid film injection holes along the circumferential direction and alternately arranging the combustion injection holes and the liquid film injection holes;
step 3.2: respectively calculating the aperture of each combustion injection hole and the aperture of each liquid film injection hole according to the flow areas of all the combustion injection holes and all the liquid film injection holes calculated in the step 2 and the number of the combustion injection holes and the liquid film injection holes preliminarily determined in the step 3.1, so as to preliminarily establish a thrust chamber model;
the specific calculation formula of the aperture of the combustion injection hole and the aperture of the liquid film injection hole is as follows:
d is the aperture of the combustion injection holes or the liquid film injection holes, and n is the number of the combustion injection holes or the liquid film injection holes;
and 4, step 4: analyzing by simulation software to obtain a final thrust chamber model;
step 4.1: simulation software carries out simulation analysis on the initially established thrust chamber model, whether the injection pressure drop of the second propellant cavity obtained through simulation is matched with the injection pressure drop of the second propellant cavity initially set in the step 2 is compared, if not, the step 2 is skipped, and the flow coefficient C is modifieddValue taking, repeating the rest steps until the injection pressure drop obtained by simulation is matched with the injection pressure drop initially set, and skipping to the step 4.2;
step 4.2: and (3) checking the spreading uniformity of the liquid film on the inner wall of the thrust chamber, if the uniformity is poor, skipping to the step 3, adjusting the number of the combustion injection holes and the number of the liquid film injection holes, and repeating the rest steps until the good spreading uniformity of the liquid film is obtained, so that the final thrust chamber model is determined.
9. The liquid film cooled rail attitude controlled engine thrust chamber of claim 8, wherein: the simulation software is ANSYS Fluent.
Priority Applications (1)
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| CN113530718A (en) * | 2021-08-31 | 2021-10-22 | 西安航天动力研究所 | Body module for hot test of rocket engine thrust chamber |
| CN114893327B (en) * | 2022-04-15 | 2023-12-26 | 西安航天动力研究所 | Method for detecting uniformity of liquid film on outer ring of pintle injector |
| CN114991997B (en) * | 2022-06-01 | 2024-09-13 | 西安航天动力研究所 | Body and thrust chamber |
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