CN103758663B - A kind of rocket based combined cycle Ejector Mode performance test motor - Google Patents
A kind of rocket based combined cycle Ejector Mode performance test motor Download PDFInfo
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- CN103758663B CN103758663B CN201410025064.7A CN201410025064A CN103758663B CN 103758663 B CN103758663 B CN 103758663B CN 201410025064 A CN201410025064 A CN 201410025064A CN 103758663 B CN103758663 B CN 103758663B
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Abstract
The invention discloses a kind of rocket based combined cycle Ejector Mode performance test motor, main rocket chamber is fixed on secondary-burning chamber front portion, intake duct is arranged on the both sides up and down of main rocket chamber respectively, inlet mouth end is fixed on main rocket chamber rear portion by support ear, intake duct rear end is connected with secondary-burning chamber flange and secondary-burning chamber by intake duct flange, and secondary-burning chamber is the straight configurations such as expansion; Main rocket chamber is connected with secondary-burning chamber coupling, adopts runner by justifying the gradual change turning side at nozzle divergence cone.Main rocket chamber adopts heptangle star and the endoporus mixed charge of adherent cast, after ignition, main rocket gas flow remains unchanged substantially, verify that rocket based combined cycle motor Ejector Mode performance is with flight Mach number and Changing Pattern highly, for runner design provides foundation.Structure of the present invention is simple, easy to maintenance, can expand the impact that research Gas Components climbs on engine performance and trajectory.
Description
Technical field
The invention belongs to rocket ramjet field, specifically, relate to a kind of rocket based combined cycle Ejector Mode performance test motor.
Background technique
Rocket based combined cycle (Rocket-Based-Combined-Cycle, RBCC) propulsion system is organically combined in same runner by rocket motor and pressed engine, optimum operation mode can be enabled under different flying heights and Mach number, give full play to rocket motor and pressed engine feature separately, make the advantage that rocket based combined cycle propulsion system also has zero-speed startup, can reuse while having high specific impulse, high thrust weight ratio.And all flight tests are by the only way which must be passed of ground research steering practical application.For rocket based combined cycle propulsion system, the application of Rocket ejector mode and Mach number Ma=0 ~ 2.5 is the keys realizing integrated runner design, the complexity of system can be effectively reduced, by flight test, the reliability of raising work, can verify that the zero-speed of rocket based combined cycle motor is taken off and autoacceleration ability.
At present, the GTX flight test project checking overall trajectory performance of the external existing U.S., as document " AffordableFlightDemonstrationoftheGTXAir-BreathingSSTOVe hicleConcept " (NASA/TM-2003-212315); The CAMUI subsonic flight test of Japan, as document " DevelopmentofRegressionFormulasforCAMUITypeHybridRockets asFunctionsofLocalO/F " (AIAAPaper2010-7117).Equally distributed three solid booster rockets are bound outward at aircraft in GTX Test item, aircraft is accelerated to 2.4 Mach, then RBCC motor starts from Asia combustion punching press Modality work, cannot verify that its zero-speed is taken off and automatic accelerating ability, and system bulky complex, test period is long, and input is large; CAMUI subsonic flight is tested, in view of the program have employed CAMUI-90P motor, thrust is 900N, and the increase of injection pipeline associated mass causes ceiling altitude to be only 500m, maximum flight Mach number is only 0.3, cannot verify that RBCC motor is in transonic speed, low supersonic flight condition injection performance.
Summary of the invention
In order to verify that the zero-speed of rocket based combined cycle motor is taken off and autoacceleration ability, reach the object that its Ejector Mode performance under Live Flying condition carries out verifying, the present invention proposes a kind of rocket based combined cycle Ejector Mode performance test motor.Object utilizes rocket based combined cycle Ejector Mode performance test motor, research is when main rocket flow realizes Secondary Control, the Ejector Mode performance of rocket based combined cycle motor is with the Changing Pattern of flying height and Mach number, the key factor that engine performance promotes is affected when finding out different Ma, the correctness of checking numerical simulation, thus provide foundation for motor runner design.
The technical solution adopted for the present invention to solve the technical problems is: comprise main rocket chamber, intake duct, secondary-burning chamber, reinforcement support ear, housing, igniter, powder charge, heat insulation layer, jet pipe, investigation on thermal protection for nozzle layer, intake duct flange, nozzle flange, secondary-burning chamber flange, main rocket chamber is positioned at the front portion of secondary-burning chamber, two intake ducts lay respectively at the both sides up and down of main rocket chamber, and symmetrical installation; Inlet mouth end is fixed on main rocket chamber rear portion by strengthening support ear, and intake duct rear end is connected with secondary-burning chamber flange and secondary-burning chamber by intake duct flange; Intake duct flange and secondary-burning chamber flange are bolted, and nozzle flange is positioned at intermediate portion on intake duct flange, and intake duct flange and secondary-burning chamber flange are divided into upper and lower two-part;
Described main rocket chamber is thin-wall barrel, front end is elliposoidal end socket, there is screw in end socket center portion, igniter one end is provided with screw thread, igniter and front end end socket are threaded connection, adopt the heptangle star of adherent cast and interior circular hole mixed charge in main rocket chamber, jet pipe and housing are bolted and install seal ring additional, adopt runner by justifying the gradual change turning side at the extending section of jet pipe; Main rocket chamber is connected with secondary-burning chamber coupling, and secondary-burning chamber is the straight configurations such as rectangular expansion, and the angle of flare is 3.03 °.
Described intake duct is S shape structure, and adopt wall thickness to be 2mm sheet material welding fabrication, four weld joints install supporting bar additional respectively.
Described main rocket chamber heat insulation layer adopts Ethylene Propylene Terpolymer material, and being 5mm with the heat insulation layer thickness of powder charge star section circumference, is 3mm with the heat insulation layer thickness of circular hole portion circumference in powder charge.
Described case material selects D406A, and thickness is 5mm.
Described investigation on thermal protection for nozzle layer material is high silica phenolic resin.
Beneficial effect
A kind of rocket based combined cycle Ejector Mode performance test motor that the present invention proposes, main rocket chamber adopts the heptangle star of adherent cast and interior circular hole mixed charge, after igniting, before powder charge burning, 2s gas flow can reach 6kg/s, and rear 10s flow is 2.5kg/s.The object regulating main rocket gas flow is reached by changing Burning area of grain.Initial time gas flow is comparatively large, and thrust is comparatively large, meets launch requirements, and after launching, gas flow reduces, and ensures that motor has high specific impulse performance.After ignition, main rocket gas flow remains unchanged, and secondary-burning chamber select expansion etc. straight determine geometric configuration, checking rocket based combined cycle motor Ejector Mode performance can be reached with flight Mach number and Changing Pattern highly, for runner design provides foundation.
Structure of the present invention is simple, easy to maintenance, also can expand the impact that research Gas Components climbs on engine performance and trajectory.
Accompanying drawing explanation
Below in conjunction with drawings and embodiments, a kind of rocket based combined cycle Ejector Mode of the present invention performance test motor is described in further detail.
Fig. 1 is rocket based combined cycle Ejector Mode performance test distribution engine schematic diagram of the present invention.
Fig. 2 is main rocket chamber structure sectional view.
Fig. 3 a, Fig. 3 b are intake duct flange, nozzle flange, secondary-burning chamber flange arrangement schematic diagram.
In figure:
1. main rocket chamber 2. intake duct 3. secondary-burning chamber 4. strengthens support ear 5. housing 6. igniter 7. powder charge 8. heat insulation layer 9. jet pipe 10. investigation on thermal protection for nozzle layer 11. intake duct flange 12. nozzle flange 13. secondary-burning chamber flange 14. screw 15. screw
Embodiment
The present embodiment is a kind of rocket based combined cycle Ejector Mode performance test motor.
Main rocket chamber adopts the heptangle star of adherent cast and the mixed charge form of interior circular hole, reach by changing Burning area of grain the object regulating main rocket gas flow, initial time gas flow is comparatively large, and thrust is larger, after launching, gas flow reduces, and ensures that motor has high specific impulse performance.After powder charge ignition, main rocket gas flow remains unchanged substantially, and secondary-burning chamber to be that expansion etc. is straight determine geometric configuration, the object of checking rocket based combined cycle motor Ejector Mode performance with the Changing Pattern of flight Mach number and height can be reached.
Consult Fig. 1, Fig. 2, Fig. 3 a, Fig. 3 b, rocket based combined cycle Ejector Mode performance test motor of the present invention is made up of main rocket chamber 1, intake duct 2, secondary-burning chamber 3, reinforcement support ear 4 and intake duct flange 11, nozzle flange 12, secondary-burning chamber flange 13.Main rocket chamber 1 is fixed on the front portion of secondary-burning chamber 3, and two intake ducts 2 are arranged on the both sides up and down of main rocket chamber 1 respectively; Intake duct 2 inlet end portion is fixed on main rocket chamber 1 rear portion by strengthening support ear 4, and intake duct 2 rear end is connected with secondary-burning chamber flange 13 and secondary-burning chamber 2 by intake duct flange 11; In the high velocity air shear action that main rocket chamber 1 exports, and extraneous and main rocket chamber 1 export the differential pressure action in high velocity, low pressure district under the air that entered by intake duct 2 injection, in secondary-burning chamber 3, export fuel-rich gas mixing with main rocket chamber 1 burn, the basis of motor generation thrust increases thrust, improves engine/motor specific impulse and reach thrust augmentation effect.Intake duct 2 is S shape structure, and lip height is 40mm, and width is 120mm, and sealing Mach number 1.4, ensures that the air that injection is come in compresses with high total pressure recovery coefficient.Intake duct 2 adopts wall thickness to be welding fabrication after the plate cutting of 2mm, and length is 300mm, longitudinal offset distance 112mm, and four weld joints install supporting bar additional respectively.S shape intake duct 2 is connected with secondary-burning chamber 3 by intake duct flange 11, strengthens support ear 4 and is fixedly connected with main rocket chamber 1, make whole intake duct 2 be equivalent to an overhang, realizes fixing reinforing function.Secondary-burning chamber 3 is the straight configurations such as expansion, and long is 0.8m, and extending section length is 0.7m, and the angle of flare is 3.03 °, and entrance opening dimension is wide × high is 120*150mm
2, outlet size is 194*150mm
2.
Main rocket chamber 1 comprises housing 5, igniter 6, powder charge 7, heat insulation layer 8, jet pipe 9 and investigation on thermal protection for nozzle layer 10.Main rocket chamber 1 is thin-wall barrel, and front end is ellipsoidal head structure, and there is screw in end socket center portion, and igniter 6 one end is provided with screw thread, and igniter 6 and front end end socket are threaded connection; Black gunpowder selected by ignition powder, and dose is 110g.Adopt the heptangle star of adherent cast and interior circular hole mixed charge in main rocket chamber 1, after igniting, before powder charge burning, 2s gas flow can reach 6kg/s, and rear 10s flow is 2.5kg/s.Jet pipe 9 and housing 5 are bolted and install seal ring additional, and jet pipe throat lining protection material is carbon/carbon compound material; Main rocket chamber 1 is connected with secondary-burning chamber 2 coupling, adopts runner by justifying the gradual change turning side at the extending section of jet pipe.
Housing 5 material is D406A, and thickness is 5mm.
Main rocket chamber 1 heat insulation layer adopts Ethylene Propylene Terpolymer material, and being 5mm with heat insulation layer 8 thickness of powder charge 7 star section circumference, is 3mm with heat insulation layer 8 thickness of circular hole portion circumference in powder charge 7.
Investigation on thermal protection for nozzle layer 10 material is high silica phenolic resin, comprising converging portion heat insulation layer, extending section heat insulation layer and larynx lining section back wall heat insulation layer.
As shown in Figure 3 a, 3 b, be intake duct flange 11 of the present invention, nozzle flange 12, secondary-burning chamber flange 13 structural representation.Intake duct flange 11, secondary-burning chamber flange 13 are square frame-shaped structure, four edges frame is provided with symmetrical screw, the two end part frame of nozzle flange 12 is provided with symmetrical screw 14, and coordinates with intake duct flange 11 and secondary-burning chamber flange frame screw; Intake duct flange 11 and secondary-burning chamber flange 13 are bolted, and nozzle flange 12 is fixed on intermediate portion on intake duct flange and is connected by screw, and intake duct flange 11 and secondary-burning chamber flange 13 are divided into upper and lower two-part; Intake duct flange 11 and all add sealing filler ring between secondary-burning chamber flange 13 and nozzle flange 12.
See table 1, calculate engine performance under typical condition by CFD, Mach number is higher, and injection when engine/motor specific impulse is higher.
Table 1
| 0Ma | 0.8Ma | 1.2Ma | |
| Highly (km) | 0 | 1.5 | 3.0 |
| Inject ratio | 0.126 | 1.0337 | 1.2317 |
| Thrust (N) | 13350 | 5802 | 6060 |
| Specific impulse (m/s) | 2384 | 2417 | 2638 |
| Thrust augmentation ratio | 5.11% | 16.8% | 28.2% |
Claims (5)
1. a rocket based combined cycle Ejector Mode performance test motor, comprise main rocket chamber, intake duct, secondary-burning chamber, housing, igniter, powder charge, intake duct flange, nozzle flange, secondary-burning chamber flange, main rocket chamber is positioned at the front portion of secondary-burning chamber, intake duct rear end is connected with secondary-burning chamber flange and secondary-burning chamber by intake duct flange, intake duct flange and secondary-burning chamber flange are bolted, main rocket chamber is connected with secondary-burning chamber coupling, it is characterized in that: also comprise reinforcement support ear, heat insulation layer, jet pipe, investigation on thermal protection for nozzle layer, two intake ducts lay respectively at the both sides up and down of main rocket chamber, and symmetrical installation, inlet mouth end is fixed on main rocket chamber rear portion by strengthening support ear, and nozzle flange is positioned at intermediate portion on intake duct flange, and intake duct flange and secondary-burning chamber flange are divided into upper and lower two-part, described main rocket chamber is thin-wall barrel, front end is elliposoidal end socket, there is screw in end socket center portion, igniter one end is provided with screw thread, igniter and front end end socket are threaded connection, adopt the heptangle star of adherent cast and interior circular hole mixed charge in main rocket chamber, jet pipe and housing are bolted and install seal ring additional, adopt runner by justifying the gradual change turning side at the extending section of jet pipe, secondary-burning chamber is the straight configurations such as rectangular expansion, and the angle of flare is 3.03 °.
2. rocket based combined cycle Ejector Mode performance test motor according to claim 1, is characterized in that: described intake duct is S shape structure, and adopt wall thickness to be 2mm sheet material welding fabrication, four weld joints install supporting bar additional respectively.
3. rocket based combined cycle Ejector Mode performance test motor according to claim 1, it is characterized in that: described main rocket chamber heat insulation layer adopts Ethylene Propylene Terpolymer material, being 5mm with the heat insulation layer thickness of powder charge star section circumference, is 3mm with the heat insulation layer thickness of circular hole portion circumference in powder charge.
4. rocket based combined cycle Ejector Mode performance test motor according to claim 1, is characterized in that: described case material selects D406A, and thickness is 5mm.
5. rocket based combined cycle Ejector Mode performance test motor according to claim 1, is characterized in that: described investigation on thermal protection for nozzle layer material is high silica phenolic resin.
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Families Citing this family (9)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN105157947B (en) * | 2015-08-19 | 2016-08-24 | 南京航空航天大学 | A kind of combination in series power air intake duct MODAL TRANSFORMATION OF A test method |
| CN105201689A (en) * | 2015-09-09 | 2015-12-30 | 西北工业大学 | Variable-geometry RBCC (rocket based combined cycle) engine for ground experiment |
| CN105351113B (en) * | 2015-11-30 | 2017-04-12 | 清华大学 | Rocket based combined engine |
| CN109101765B (en) * | 2018-09-19 | 2022-06-14 | 西北工业大学 | Mechanism modeling method for large-package-width fast-domain propulsion system of combined power aircraft |
| CN109738196B (en) * | 2019-01-18 | 2020-07-31 | 中国人民解放军国防科技大学 | Ramjet performance space-ground conversion method based on ground direct connection test |
| CN110566365B (en) * | 2019-09-29 | 2021-01-08 | 中国人民解放军国防科技大学 | Mode-switchable solid combined engine and missile |
| CN111594344A (en) * | 2020-05-01 | 2020-08-28 | 西北工业大学 | Small-scale two-stage rocket combined ramjet engine |
| CN113090416B (en) * | 2021-04-27 | 2022-02-22 | 西北工业大学 | Simulation experiment device for rocket stamping combined air inlet channel |
| CN114352437A (en) * | 2022-01-07 | 2022-04-15 | 北京理工大学 | Solid fuel stamping combined engine suitable for wide Mach number flight |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| US4277940A (en) * | 1979-07-25 | 1981-07-14 | United Technologies Corporation | Integral rocket-ramjet closed loop fuel control system |
| JPS5681246A (en) * | 1979-12-03 | 1981-07-03 | Nissan Motor Co Ltd | Ram rocket |
| CN102400814B (en) * | 2011-10-27 | 2013-09-18 | 北京航空航天大学 | Solid-liquid hybrid rocket ramjet for test |
| CN103244309B (en) * | 2013-04-28 | 2016-01-20 | 湖北航天技术研究院总体设计所 | A kind of solid-rocket punching press combined engine housing |
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