WO2016127813A1 - Moteur de statoréacteur à combustion supersonique de grande taille et chambre de combustion de section pétaliforme tridimensionnelle associée - Google Patents

Moteur de statoréacteur à combustion supersonique de grande taille et chambre de combustion de section pétaliforme tridimensionnelle associée Download PDF

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Publication number
WO2016127813A1
WO2016127813A1 PCT/CN2016/072424 CN2016072424W WO2016127813A1 WO 2016127813 A1 WO2016127813 A1 WO 2016127813A1 CN 2016072424 W CN2016072424 W CN 2016072424W WO 2016127813 A1 WO2016127813 A1 WO 2016127813A1
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WIPO (PCT)
Prior art keywords
section
combustion chamber
petal
scramjet engine
fuel injection
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PCT/CN2016/072424
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English (en)
Chinese (zh)
Inventor
孙明波
王振国
赵玉新
赵国焱
梁剑寒
谭建国
Original Assignee
中国人民解放军国防科学技术大学
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Publication of WO2016127813A1 publication Critical patent/WO2016127813A1/fr

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers

Definitions

  • the present invention relates to a large scale scramjet engine and a combustion chamber therefor.
  • the gas temperature is high ( ⁇ 2650K)
  • the pressure is high (3 ⁇ 5atm)
  • the heat flow is large
  • the flow rate of cooling fuel is small, making the combustion chamber heat protection extremely difficult.
  • the requirements for thermal protection impose strong constraints on the design of the combustion chamber.
  • the common forms are support plates, slopes, cantilever beams, delta wings and the like.
  • This type of injection directly injects fuel into the main stream of supersonic flow, providing good fuel spatial distribution and mixing effects, and the recirculation zone formed at the bottom of the injection device can provide a stable flame effect.
  • the paper "Pylon Fuel Injector Design for a Scramjet Combustor” (AIAA paper 2007-5404) describes a support plate configuration for large-scale combustion chambers. The support plates are perpendicular to the direction of flow and are staggered with trailing edges for fuel and Mix of air.
  • plug-in plates, delta wings or ramp fuel injection in the combustion chamber can solve the problem of fuel jet penetration and mixing, because these plug-in devices are directly exposed to the high total temperature supersonic flow, With high heat load, it is difficult to solve the thermal protection problem at present.
  • an aspect of the invention provides a three-dimensional petal-shaped cross-section combustion chamber of a scramjet engine, comprising: a fuel injection section, a flame stabilization section and an expansion section, wherein the wall surface of the fuel injection section is convex and concave a lobe structure alternately formed in the circumferential direction, the lobes of the lobed structure gradually approach the central axis along the flow direction, and the lobed structure continues along the flow direction to the fuel injection section
  • a petal-like structure is formed at the end, and a second group of nozzle holes is disposed at a vertex position of each of the protrusions at the end of the combustion jetting section, and a first group of nozzle holes is disposed at a vertex position of each of the recesses upstream of the second group of the nozzle holes.
  • the first group of nozzle holes are located on a section of the starting position of the protrusion.
  • the fuel injection section end position is provided with a cavity flame stabilizer for forming a back pressure recirculation zone.
  • the above-described cavity flame stabilizer comprises a cylindrical body and an end wall for closing a gap between the cylindrical body and the end of the combustion jetting section.
  • cross section of the flame stabilizing section at the rear end of the cavity flame stabilizer is a petal-like structure that engages with the end of the combustion jetting section.
  • the petal-like structure at the end of the fuel injection section extends in the flow direction over the entire length of the flame stabilizing section.
  • the above-mentioned lobe structure is a center-symmetric four-lobed structure.
  • the flow channel wall surface of the fuel injection section is a wave-eliminating wall surface designed by wave elimination.
  • the above-mentioned wave-eliminating wall surface is obtained by bidirectional streamline tracking and streamline fusion.
  • a large scale scramjet engine comprising a combustion chamber, characterized in that the combustion chamber is a three-dimensional petal-section combustion chamber according to the scramjet engine described above.
  • the lobed structure can divide the flowing air into a plurality of air streams in the future, which facilitates the injection mixing of the fuel, and provides the injection hole b at the petal protrusion of the fuel injection section 1 to increase the injection.
  • the penetration depth of the fuel, the combination of the convex injection of the petal combustion chamber and the concave injection, solves the uniform mixing of the fuel.
  • there is no plug-in plate in the three-dimensional petal-shaped cross-section combustion chamber and the thermal protection constraint (whether active cooling or passive thermal protection) can be realized in the design of the combustion chamber wall.
  • FIG. 1 is a perspective structural view of a first embodiment of a three-dimensional petal-section combustion chamber of a scramjet engine according to the present invention
  • Figure 2 is a schematic plan view of the combustion chamber of Figure 1;
  • Figure 3 is a left side view of the combustion chamber shown in Figure 2;
  • Figure 4 is a cross-sectional view taken along line A-A of the combustion chamber shown in Figure 3;
  • Figure 5 is a cross-sectional view taken along line B-B of the combustion chamber shown in Figure 3;
  • Figure 6 is an enlarged plan view showing a portion C of the combustion chamber shown in Figure 4.
  • Figure 7 is an enlarged plan view showing a portion D of the combustion chamber shown in Figure 5;
  • Figure 8 is a schematic end view of the cross section of the combustion chamber E-E shown in Figure 2;
  • Figure 9 is a schematic end view of the cross section of the combustion chamber F-F shown in Figure 2;
  • Figure 10 is a schematic end view of the cross section of the combustion chamber G-G shown in Figure 2;
  • Figure 11 is a schematic end view of the cross section of the combustion chamber H-H shown in Figure 2;
  • Figure 12 is a perspective structural view of a second embodiment of a three-dimensional petal-section combustion chamber of a scramjet engine in accordance with the present invention.
  • Figure 13 is a schematic plan view showing the structure of the combustion chamber shown in Figure 12;
  • Figure 14 is a left side view of the combustion chamber shown in Figure 13;
  • Figure 15 is a cross-sectional view taken along line A-A of the combustion chamber shown in Figure 14;
  • Figure 16 is a cross-sectional view taken along line B-B of the combustion chamber shown in Figure 14;
  • Figure 17 is a schematic side view of the cross section of the combustion chamber E-E shown in Figure 13;
  • Figure 18 is a schematic end view of the cross section of the combustion chamber F-F shown in Figure 13;
  • Figure 19 is a schematic end view of the cross section of the combustion chamber G-G shown in Figure 13;
  • Figure 20 is a schematic side view of the cross section of the combustion chamber H-H shown in Figure 13;
  • Figure 21 is an end view showing a section of the combustion chamber I-I shown in Figure 13;
  • FIGS. 1 through 11 illustrate the shape and configuration of a three-dimensional petal-section combustion chamber in accordance with the present invention.
  • the shape and structure of the three-dimensional petal-section combustion chamber of the present invention includes a fuel injection section 1, a flame stabilization section 2, and an expansion section 3.
  • the wall surface of the fuel injection section is a lobed structure in which the protrusions 12 and the recesses 13 are alternately arranged in the circumferential direction, wherein the lobed structure gradually develops along the flow direction on the basis of ensuring the area expansion rate, in the fuel
  • a petal-shaped cross section is formed near the outlet of the injection section and at the position of the flame stabilizing section 2, and a second set of injection holes 15 is provided at the petal protrusion 12 of the fuel injection section 1, and is disposed upstream of the fuel injection section 1
  • the first set of orifices 14 form a combined jet.
  • the wall surface which protrudes inward with respect to the circular inlet of the combustion chamber is referred to as a projection, and the connection of the adjacent two projections is referred to as a recess.
  • the lobing structure can divide the flowing air into a plurality of air streams in the future, which facilitates the injection mixing of the fuel, and provides a second group of nozzle holes 15 at the petal protrusion of the fuel injection section 1. Increase the penetration depth of the injected fuel, and combine the convex injection of the petal combustion chamber with the concave injection to solve the uniform mixing of the fuel.
  • the wall surface of the combustion chamber of the invention is a curved wall surface with continuous variation and low curvature, and the thermal protection constraint (whether active cooling or passive thermal protection) can be well realized in the design of the combustion chamber wall.
  • Figure 2 is a schematic plan view of the combustion chamber of Figure 1
  • Figure 3 is a left side view of the combustion chamber of Figure 2, as shown in Figure 3, the inlet 11 of the fuel injection section 1 of the combustion chamber is circular, wherein The four projections 12 extend a certain distance inside the circle.
  • Figure 4 is a cross-sectional view taken along line AA of the combustion chamber of Figure 3, i.e., cut at the apex of the projection 12, the projection 12 being formed near the inlet of the fuel injection section 1, gradually approaching the center in the flow direction, and to the fuel injection
  • the depth of penetration of the projection 12 at the end of the segment reaches a maximum, after which the depth of penetration is maintained in the flame stabilizing section, and then gradually moves away from the central axis in the flow direction until the projection disappears.
  • Figure 5 is a cross-sectional view taken along the line BB of the combustion chamber shown in Figure 3, i.e., cut at the apex of the recess 13 of the recess 13, as can be seen in conjunction with Figures 4 and 5, the fuel injection section 1 has two sets of spray holes, the first set of sprays The hole 14 is located at the apex position of the recess in the section where the starting position of the projection 12 is located, and the second group of the nozzle holes 15 is located at a position where the projection 12 projects to the maximum depth.
  • Figure 6 is an enlarged plan view showing a portion C of the combustion chamber shown in Figure 4
  • Figure 7 is an enlarged view of a portion D of the combustion chamber shown in Figure 5.
  • the first set of orifices 14 are located at the apex of the recess 13
  • the second set of orifices 15 are located at the apex of the projections 12.
  • FIG. 8 is a schematic view of the end face of the combustion chamber E-E of Figure 2, as shown in Figure 8, the cut-out is located at the position of the first set of orifices, i.e., near the starting point of the projection 12.
  • 9 is a schematic end view of the cross section of the combustion chamber FF shown in FIG. 2, as shown in FIG. 9, the cut position is located at the position where the second set of injection holes are located, that is, the position where the protrusion 12 reaches the maximum penetration depth
  • FIG. 10 is a schematic view of the end face at the cross section of the combustion chamber GG shown in FIG. 2, as shown in FIG. 10, which is located in the flame stabilizing section.
  • Figure 11 is a schematic view of the end face at the section of the combustion chamber H-H shown in Fig. 2. As shown in Fig. 11, the section is located at the position of the expansion section, wherein the projection gradually gathers toward the circular outlet.
  • the lobe structure is a four lobe structure. In one other embodiment, it is a three-lobed structure, in another embodiment a five-lobed structure, or another number of lobes.
  • the flame stabilizing section is a petal-like structure, and the lobe structure of the fuel injection section 1 is formed along the flow direction on the basis of ensuring the area expansion ratio.
  • the end of the expansion section has a circular cross section for engaging with the tail nozzle, and the petal-like structure of the flame stabilization section gradually gathers when it develops to the expansion section, and finally has a circular section.
  • the flow path profile of the combustion chamber is a wave-eliminating wall that has been subjected to wave-breaking design and/or experimental modification to reduce or eliminate shock loss.
  • the flow path profile of the burner of the present invention is obtained by bidirectional streamline tracking and streamline fusion.
  • Fig. 22 to Fig. 24 for a wide range of Mach numbers, there is no obvious shock wave structure in the flow field of the petal-shaped section combustion chamber, and the internal flow loss is small.
  • the following is a comparison table of the total pressure loss of the three-dimensional petal-section combustion chamber of the flow path profile obtained by the two-way flow line tracking and the streamline fusion method and the circular expansion chamber of the same area expansion ratio circular section.
  • the first embodiment is identical in structure and shape to the first embodiment in the fuel injection section, and the second embodiment is different from the first embodiment in that it is provided at the end of the fuel injection section.
  • Cavity flame stabilizer 4 is identical to the first embodiment in the fuel injection section, and the second embodiment is different from the first embodiment in that it is provided at the end of the fuel injection section.
  • the cavity flame stabilizer 4 is formed in the combustion chamber for the low Mach number of the hypersonic flight.
  • the leeward recirculation zone is formed to achieve flame stabilization, and the combustion chamber flame stability under low total temperature conditions is achieved.
  • the cavity flame stabilizer 4 includes a cylindrical body 41 at the end of the fuel injection section 1, and the outer diameter of the cylindrical body 41 is substantially or slightly larger than the outer end of the end of the fuel injection section 1. The diameter of the circle, and the gap between the fuel injection section 1 and the cylindrical body 41 is closed by the end wall 42, thereby forming the cavity flame stabilizer 4.
  • the EE section is the position of the first group of nozzles, as shown in Fig. 18, the FF section is the position of the second group of nozzles, as shown in Fig. 19, the GG section is located in the cavity flame stabilizer. 4, it can be seen from Fig. 19 that the position of the cavity flame stabilizer 4 does not form a petal-like structure.
  • the H-H section is located in the flame stabilizing section, and the wall surface of the flame stabilizing section is a petal-like structure which is the same as the petal-like structure at the end of the fuel injection section.
  • the I-I section is in the expanded section, in which the projections gradually gather toward the circular outlet.
  • the three-dimensional petal-shaped cross section has the great advantage of fuel injection and mixing at the same time of low loss, and the uniform injection on a short distance can be realized by the combination injection of the appropriate positions of the protrusion and the recess;
  • the leeward recirculation zone in the petal bulge is larger than the recirculation zone of the device such as the support plate, which produces a better flame stability effect in the core flow zone;
  • three-dimensional petal-shaped curved wall surface, low curvature and smooth transition can be used for existing active cooling thermal protection, passive thermal protection technology
  • the flow path surface transformation formed by the rotary characteristic line method, the bidirectional flow line tracking and the streamline fusion method has no obvious shock wave structure in the flow field in the wide inlet Mach number range, and the internal flow loss is small.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)
  • Nozzles For Spraying Of Liquid Fuel (AREA)

Abstract

L'invention concerne un moteur de statoréacteur à combustion supersonique, de grande taille, et une chambre de combustion de section pétaliforme tridimensionnelle associée. La chambre de combustion de section pétaliforme tridimensionnelle comprend une section (1) d'injection de carburant, une section (2) de stabilisation de flamme et une section (3) de détente. La face de mur de la section (1) d'injection de carburant présente une structure en lobes dans laquelle des bosses (12) et des évidements (13) sont formés en alternance dans la direction de la circonférence. Les bosses (12) de la structure en lobes s'approchent progressivement d'un axe central le long d'une direction d'écoulement. La structure en lobes s'étend jusqu'à l'extrémité arrière de la section (1) d'injection de carburant, formant ainsi une structure pétaliforme. Un second groupe d'orifices (15) sont formés au niveau de la position de sommet de chacune des bosses (12) à l'extrémité arrière de la section (1) d'injection de carburant. Un premier groupe d'orifices (14) sont formés au niveau de la position de sommet de chacun des évidements (13) en amont du second groupe d'orifices (15), formant ainsi une injection combinée.
PCT/CN2016/072424 2015-02-13 2016-02-04 Moteur de statoréacteur à combustion supersonique de grande taille et chambre de combustion de section pétaliforme tridimensionnelle associée WO2016127813A1 (fr)

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CN201510079040.4 2015-02-13
CN201510079040.4A CN104654362B (zh) 2015-02-13 2015-02-13 大尺度超燃冲压发动机和三维花瓣形截面燃烧室

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CN112361379A (zh) * 2020-11-18 2021-02-12 中国人民解放军国防科技大学 一种超声速凹腔燃烧室的点火结构及超燃冲压发动机
CN112668201A (zh) * 2021-01-11 2021-04-16 中国人民解放军国防科技大学 基于进气道非均匀气流的超声速燃烧室燃油喷注设计方法

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CN104654362B (zh) * 2015-02-13 2016-08-24 中国人民解放军国防科学技术大学 大尺度超燃冲压发动机和三维花瓣形截面燃烧室
CN105180212B (zh) * 2015-09-02 2017-06-16 中国人民解放军国防科学技术大学 超燃冲压发动机燃烧室
CN105180211B (zh) * 2015-09-02 2017-06-16 中国人民解放军国防科学技术大学 具有凹腔火焰稳定器的燃烧室及超燃冲压发动机
CN105423342B (zh) * 2016-01-12 2018-08-21 西北工业大学 微型发动机燃烧室凹腔壁面蒸发管
CN109139267B (zh) * 2018-09-11 2019-10-29 中国人民解放军国防科技大学 超声速流动增混装置
CN109519284A (zh) * 2018-12-12 2019-03-26 北京动力机械研究所 一种燃烧室隔热屏
CN112014420B (zh) * 2020-09-03 2022-02-18 中南大学 加热设备
CN112524642B (zh) * 2020-12-04 2022-08-23 中国人民解放军国防科技大学 一种大尺度冲压发动机燃烧室及冲压发动机
CN114623467B (zh) * 2022-01-27 2023-05-16 北京盈天航空动力科技有限公司 一种微小型涡喷发动机波瓣型火焰筒结构
CN115419917A (zh) * 2022-07-29 2022-12-02 西安航天动力研究所 一体式异质多相流掺混稳焰装置及组合动力发动机燃烧室

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CN112361379B (zh) * 2020-11-18 2022-03-18 中国人民解放军国防科技大学 一种超声速凹腔燃烧室的点火结构及超燃冲压发动机
CN112668201A (zh) * 2021-01-11 2021-04-16 中国人民解放军国防科技大学 基于进气道非均匀气流的超声速燃烧室燃油喷注设计方法

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