CN109250149A - Flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation - Google Patents
Flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation Download PDFInfo
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- CN109250149A CN109250149A CN201811119225.3A CN201811119225A CN109250149A CN 109250149 A CN109250149 A CN 109250149A CN 201811119225 A CN201811119225 A CN 201811119225A CN 109250149 A CN109250149 A CN 109250149A
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
- B64F5/60—Testing or inspecting aircraft components or systems
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- Aviation & Aerospace Engineering (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention discloses the flow tunnel testing devices for air suction type hypersonic vehicle radome fairing separation simulation, including model equipment and jet apparatus.The device is based on integrated design thought, experimental rig has been comprehensively considered to model support, jet flow gas supply, separating distance adjustment requirement similar with interstage area shape is guaranteed, aircraft precursor after designed correction of the flank shape has the function of model support, jet flow gas supply and separating distance adjustment, and it ensure that the similitude of interstage area shape, designed gas supply adapting rod has the function of model support and jet flow gas supply, a whole set of experimental rig is easy to assemble and disassemble, easy to use.Flow tunnel testing device of the invention solves the model support that current test faces, separating distance adjustment, counter pushes away the key technical problems such as hot jet simulation, it ensure that the similitude of radome fairing, be not introduced into it is additional be difficult to modified interference, obtain and counter push away reliable radome fairing aerodynamic loading data under jet flow and incoming flow interaction.
Description
Technical field
The invention belongs to hypersonic wind tunnel experiment technical fields, and in particular to be used for air suction type hypersonic vehicle radome fairing
The flow tunnel testing device of separation simulation.
Background technique
Air-breathing hypersonic vehicle uses scramjet engine for power, can do hypersonic fly in endoatmosphere
Row.But scramjet engine can only start under the conditions of hypersonic, the competence exertion optimality under design point flying condition
Energy.Therefore, rocket need to be used Air-breathing hypersonic vehicle boosting to design point flying condition.
During boosting of certain Air-breathing hypersonic vehicle before reaching design point flying condition, using body of revolution
Radome fairing protects the precursor, air intake duct and its relevant device of the aircraft, from the broken of the harsh environments such as aerodynamic force, Aerodynamic Heating
It is bad.It is power in endoatmosphere entirety high-speed separation using small-sized retro-rocket, it may be assumed that open first after design point flying condition has
Three retro-rocket engines being evenly distributed on dynamic radome fairing, radome fairing is extracted forward from aircraft, wait separate one
Retro-rocket engine is closed after set a distance, a side for restarting side-lower on radome fairing pushes away rocket engine, by radome fairing
Toward laterally casting aside, to complete the overall process of radome fairing separation.The very small boat of dynamic pressure is separated under altitude air lean environment
Its separation is different, and the dynamic pressure that radome fairing pulls out cover separation in dense atmosphere is up to tens kPa, the aerodynamic characteristic of radome fairing and quiet
Stability characteristic seriously affects the interaction for pushing away rocket jet and incoming flow by small size back.It was separated by wind tunnel test acquisition
The aerodynamic characteristic of radome fairing and static stability characteristic in journey, to separation emulation and separation scheme great significance for design.
In the hypersonic stage separation simulation test carried out in the past, support sting support first-level model is generallyd use, is adopted
Second-level model is supported with abdomen supporting plate.Due to the influence of abdomen supporting plate interference, so that the aerodynamic loading of level-one, second-level model occurs partially
It moves, level-one, the normal force coefficient of second-level model and pitching moment coefficient all should be 0 when the zero degree angle of attack, but the interference of abdomen supporting plate makes
The two coefficients are obtained all to be not zero." zero passage " modification method is generally used in test, i.e., by the normal force coefficient under the zero degree angle of attack
Increment caused by being interfered with pitching moment coefficient as abdomen supporting plate, for correcting the normal force coefficient and pitching power under other angles of attack
Moment coefficient, but there is a degree of error in itself under non-zero-incidence in invariable correction amount, and applicability is limited.If
Conventional support mode is used in the test of certain Air-breathing hypersonic vehicle radome fairing separation simulation, the anti-lateral jet that pushes away will
It intercouples, cannot be distinguished with abdomen supporting plate interference effect, so that the influence amount of abdomen supporting plate interference is difficult to correct or deduct.
Summary of the invention
Technical problem to be solved by the invention is to provide the wind for air suction type hypersonic vehicle radome fairing separation simulation
Hole experimental rig.
Flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention, its main feature is that: institute
The experimental rig stated includes model equipment and jet apparatus；
The model equipment includes radome fairing head, radome fairing tail portion, through-hole cover board, six component ring type balances, gas supply switching
The aircraft fuselage after aircraft precursor, correction of the flank shape, inner flow passage baffle, support sting and balance clamp nut after bar, correction of the flank shape；Institute
The inner flow passage baffle and support sting stated are integrally machined integral, and the aircraft fuselage after correction of the flank shape is mounted on inner flow passage baffle
On, the aircraft precursor after correction of the flank shape is mounted on the aircraft fuselage after correction of the flank shape, and gas supply adapting rod is mounted on the flight after correction of the flank shape
On device precursor, six component ring type balances are mounted on gas supply adapting rod, and radome fairing tail portion is mounted on six component ring type balances, whole
Stream capouch portion is mounted on radome fairing tail portion, and through-hole cover board is mounted on radome fairing head, constitutes model equipment；
The jet apparatus includes that retro-nozzle, gas supply adapting rod, the aircraft precursor after correction of the flank shape, red copper pipeline and screw thread are anti-
To nut；Red copper pipeline is mounted on the rear end of the aircraft precursor after correction of the flank shape, and gas supply adapting rod is mounted on the aircraft after correction of the flank shape
The front end of precursor, retro-nozzle are mounted on gas supply adapting rod by the reversed nut of screw thread, and each contact surface is all made of red copper pad
Circle sealing, constitutes jet apparatus；Red copper pipeline is connected to extraneous gas source, and the normal temperature compressed air of extraneous gas source is imported into correction of the flank shape
Aircraft precursor afterwards is supplied adapting rod and enters retro-nozzle.
Have two groups of mounting-positioning holes on the ring flange of the described gas supply switching rod rear end, the uniformly distributed interior hexagonal counterbore of inner ring and
Pin hole is used to for six component ring type balances to be fixed on gas supply adapting rod, and the uniformly distributed interior hexagonal counterbore in outer ring and pin hole are used for
Gas supply adapting rod is fixed on the aircraft precursor after correction of the flank shape.
Aircraft precursor after the correction of the flank shape has 4, the separating distance that the aircraft precursor after each correction of the flank shape is adjusted point
Not Wei 0.0D, 0.8D, 1.6D and 2.2D, D is aircraft fuselage diameter, has inside the aircraft precursor after each correction of the flank shape gas supply logical
Hole, one end and red copper pipeline connection for vent hole, for the other end gas supply adapting rod connection of vent hole.
The conical nozzle of the retro-nozzle is chamfer, and the angle of beveling is identical as the cone angle on radome fairing head；It is counter to push away spray
The short side exit cross-section of pipe and the relationship of nozzle throat cross section meet that jet flow pressure ratio is equal and the similar simulation of momentum of the jet
Criterion.
The through-hole cover board allows retro-nozzle to stretch out inside model, and through-hole cover board is not contacted with retro-nozzle, stayed
There is the gap of about 0.7mm, the anti-direct Thrust for pushing away jet flow is avoided to act on six component ring type balances.
The inside of the six components ring type balance is hollow structure, allows to supply that adapting rod is contactless passes through.
Aircraft fuselage after the correction of the flank shape clips the later half of aircraft fuselage from the appropriate location of inner flow passage expansion segment
Portion, retained fuselage sections guarantee that inner flow passage and shape are constant, with short form test model；To place red copper pipeline, fuselage upper half
Portion's fluting；To mitigate model weight, back body machined lightening hole.
Diversion trench and deflection cone are arranged on the inner flow passage baffle, air-flow is exported from aircraft inner flow passage,
Diversion trench discharge area is sufficiently large, it is ensured that inner flow passage baffle does not influence the venting capability of aircraft inner flow passage.
The retro-nozzle is the core component of jet apparatus, for simulating the gaseous jet of retro-rocket engine,
The conical nozzle chamfer comprising 3,1 shared to stay room and pressure monitoring hole, and pressure monitoring hole is used when jet pressures are debugged
In installation pressure sensor, stay whether chamber pressure meets the requirements with judgement；Between retro-nozzle and gas supply adapting rod, using small convex
It rises and little groove is done circumferentially positioned, tensed using the reversed nut of screw thread.
The conical nozzle of the retro-nozzle, on the basis of retro-rocket engine export shape and gaseous jet parameter
On, foundation jet flow pressure ratio is equal and momentum of the jet analog simulation criterion designs.Because jet pipe is chamfer, it is exported in conical nozzle short side
Region afterwards, jet parameters are uneven on the cross section perpendicular to nozzle axis.When Nozzle Design, conical nozzle short side is gone out
Theory outlet of the cross section as jet pipe at mouthful is the anti-input ginseng for pushing away hot jet simulation with the jet parameters on the cross section
Number.
Flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention in test basis
4 aircraft precursors are selected in the requirement of separating distance, are played 4 different separating distances of adjustment, are kept interstage section shape and cold
The effect of jet flow gas supply.
Flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention has the advantage that
1. support part is entirely located in the inside of radome fairing, it is not destroyed the integrality of radome fairing, counter to push away jet flow mutual with incoming flow
The interference flowing field of effect is not destroyed by other disturbing factors.
2. the precursor configuration near aircraft inlet lip is retained substantially, the stream near aircraft lip is not destroyed
, the venting capability of aircraft inner flow passage is not influenced.
3. realizing guarantor's type adjustment of separating distance under each separating distance, interstage area flight is all maintained as far as possible
The similitude of device shape reduces the distortion factor of simulation.
Flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention is based on integration
The model and jet apparatus of thought design solve model support, separating distance adjustment that current test faces, counter push away jet flow mould
Quasi- equal key technical problems, ensure that the similitude of radome fairing, are not introduced into the additional modified interference that is difficult to, and obtain and counter push away spray
Stream and reliable radome fairing aerodynamic loading data under incoming flow interaction.
Detailed description of the invention
Fig. 1 is the structure of the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Schematic diagram (main view)；
Fig. 2 is the structural representation of the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Scheme (cross-sectional view)；
Fig. 3 is the model dress in the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Set partial enlarged view；
Fig. 4 is the jet flow dress in the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Set partial enlarged view；
After Fig. 5 is the correction of the flank shape in the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Aircraft precursor schematic diagram (0.0D)；
After Fig. 6 is the correction of the flank shape in the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Aircraft precursor schematic diagram (0.8D)；
After Fig. 7 is the correction of the flank shape in the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Aircraft precursor schematic diagram (1.6D)；
After Fig. 8 is the correction of the flank shape in the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Aircraft precursor schematic diagram (2.2D)；
Fig. 9 is the nozzle with scarfed exit plane in the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Partial enlarged view；
In figure, 1. radome fairing head, 2. radome fairing tail portion, 3. through-hole cover board, 4. 6 component ring type balance 5. supplies adapting rod
6. 8. inner flow passage baffle of aircraft fuselage, 9. support sting, 10. balance after 7. correction of the flank shape of aircraft precursor after correction of the flank shape compresses
11. retro-nozzle of nut, 12. red copper pipeline, the 13. reversed nut of screw thread.
Specific embodiment
The present invention is described in detail below with reference to the accompanying drawings and embodiments.
As shown in Fig. 1 ~ 4, the flow tunnel testing device for air suction type hypersonic vehicle radome fairing separation simulation of the invention
Including model equipment and jet apparatus；
The model equipment includes radome fairing head 1, radome fairing tail portion 2, through-hole cover board 3, six component ring type balances 4, gas supply
Aircraft fuselage 7, inner flow passage baffle 8, support sting 9 and the day concora crush after aircraft precursor 6, correction of the flank shape after adapting rod 5, correction of the flank shape
Tight nut 10；The inner flow passage baffle 8 and support sting 9 is integrally machined integral, and the aircraft fuselage 7 after correction of the flank shape is mounted on
On inner flow passage baffle 8, the aircraft precursor 6 after correction of the flank shape is mounted on the aircraft fuselage 7 after correction of the flank shape, and gas supply adapting rod 5 is pacified
On aircraft precursor 6 after correction of the flank shape, six component ring type balances 4 are mounted on gas supply adapting rod 5, and radome fairing tail portion 2 is installed
On six component ring type balances 4, radome fairing head 1 is mounted on radome fairing tail portion 2, and through-hole cover board 3 is mounted on radome fairing head
On, constitute model equipment；
The jet apparatus includes aircraft precursor 6,12 and of red copper pipeline after retro-nozzle 11, gas supply adapting rod 5, correction of the flank shape
The reversed nut 13 of screw thread；Red copper pipeline 12 is mounted on the rear end of the aircraft precursor 6 after correction of the flank shape, and gas supply adapting rod 5, which is mounted on, to be repaired
The front end of aircraft precursor 6 after shape, retro-nozzle 11 is mounted on gas supply adapting rod 5 by the reversed nut 13 of screw thread, each
Contact surface is all made of copper washer sealing, constitutes jet apparatus；Red copper pipeline 12 is connected to extraneous gas source, by the normal of extraneous gas source
Warm compressed air imported into the aircraft precursor 6 after correction of the flank shape, enters retro-nozzle 11 through gas supply adapting rod 5.
There are two groups of mounting-positioning holes, the uniformly distributed interior hexagonal counterbore of inner ring on the ring flange of 5 rear end of gas supply adapting rod
It is used to for six component ring type balances 4 to be fixed on gas supply adapting rod 5, outer ring uniformly distributed interior hexagonal counterbore and pin hole with pin hole
The aircraft precursor 6 being fixed on after correction of the flank shape for adapting rod 5 will to be supplied.
As shown in Fig. 5 ~ 8, the aircraft precursor 6 after the correction of the flank shape has 4, and the aircraft precursor 6 after each correction of the flank shape is adjusted
Whole separating distance is respectively 0.0D, 0.8D, 1.6D and 2.2D, and D is aircraft fuselage diameter, the aircraft precursor after each correction of the flank shape
Have inside 6 for vent hole, be connected to for one end of vent hole with red copper pipeline 12, supplies adapting rod 5 for the other end of vent hole
Connection.
As shown in figure 9, the conical nozzle of the retro-nozzle 11 is chamfer, the angle of beveling and the cone on radome fairing head 1
Angle is identical；The short side exit cross-section of retro-nozzle 11 and the relationship of nozzle throat cross section meet that jet flow pressure ratio is equal and jet flow
The similar simulation rules of momentum.
Embodiment 1
The present embodiment assembles jet apparatus first, on the aircraft fuselage after being mounted on correction of the flank shape, after jet pressures debugging to be done,
Jet apparatus and model equipment are fitted together again, constitute complete experimental rig.
1. assembling jet apparatus and jet pressures debugging, complete according to the following steps:
1a. red copper pipeline 12 passes through the centre bore of 7 front end of aircraft fuselage after correction of the flank shape, before the aircraft after the correction of the flank shape of 0.0D
Body 6 is docked with flange method, soket head cap screw fastening, copper washer sealing；
Aircraft precursor 6 after the correction of the flank shape of 1b.0.0D is cooperated with cylinder, the mode of pin positioning, the flight after being mounted on correction of the flank shape
On device fuselage 7, aircraft fuselage is mounted on wind-tunnel attack angle mechanism by inner flow passage baffle 8 and support sting 9；
1c. supplies adapting rod 5 and is mounted on repairing for 0.0D in a manner of flange docking, soket head cap screw fastening, copper washer sealing
The front end of aircraft precursor 6 after shape；
1d. retro-nozzle 11 is mounted on gas supply 5 front end of adapting rod by the reversed nut 13 of screw thread, is sealed with copper washer；
The assembling of 1e. jet apparatus finishes, and red copper pipeline 12 is connected to wind tunnel wall gas source, and room is stayed in installation monitoring on retro-nozzle
The sensor of pressure；
1f. pings gas, check supply air line and jet apparatus whether gas leakage, after leak detection, retro-nozzle 11 is stayed into chamber pressure
Power is debugged to target value, and the cold jet flow control system parameter of wind-tunnel is recorded；
1g. replaces the aircraft precursor 6 after the correction of the flank shape of 2.2D, repeats step 1a~1f, before obtaining the aircraft using the specification
When body, the cold jet flow control system parameter of the corresponding wind-tunnel of retro-nozzle goal pressure in room.
When 1h. is according to aircraft precursor after the correction of the flank shape for using 0.0D and 2.2D, the cold jet flow control system parameter of wind-tunnel, line
When property interpolation obtains the aircraft precursor after the correction of the flank shape using 0.8D and 1.6D, wind corresponding to retro-nozzle goal pressure in room
The cold jet flow control system parameter in hole.
2. fitting together jet apparatus and model equipment after jet pressures are debugged, experimental rig is constituted, is carried out
Wind tunnel test is completed according to the following steps:
2a. is mounted on integral inner flow passage baffle 8 and support sting 9 is processed on the attack angle mechanism of wind-tunnel；
Aircraft fuselage 7 after 2b. correction of the flank shape is mounted on inner flow passage baffle 8 in a manner of soket head cap screw fastening；
2c. red copper pipeline 12 passes through the centre bore of 7 front end of aircraft fuselage after correction of the flank shape, before the aircraft after the correction of the flank shape of 0.0D
Body 6 is docked with flange method, soket head cap screw fastening, copper washer sealing；
Aircraft precursor 6 after the correction of the flank shape of 2d.0.0D is cooperated with cylinder, the mode of pin positioning, the flight after being mounted on correction of the flank shape
On device fuselage 7；
Six component ring type balance 4 of 2e. is docked by flange, soket head cap screw fastens, gas supply switching is mounted in a manner of pin positioning
The rear end of bar 5, gas supply adapting rod 5 are passed through from the inside of six component ring type balances 4；
2f. supplies adapting rod 5 in its back-end with the side of flange docking, soket head cap screw fastening, pin positioning, copper washer sealing
Formula is mounted on the front end of the aircraft precursor 6 after the correction of the flank shape of 0.0D；
2g. radome fairing tail portion 2 is sleeved on the front end of six component ring type balances 4, cone match, flat key positioning, balance clamp nut
10 fastenings；
2h. retro-nozzle 11 is mounted on gas supply 5 front end of adapting rod by the reversed nut 13 of screw thread, is sealed with copper washer, uses
Protrusion and groove positioning；
2i. radome fairing head 1 is cooperated with cylinder, and the mode of pin positioning is mounted on radome fairing tail portion 2, and through-hole cover board 3 uses
Screw is mounted on radome fairing head 1；
2j. carries out wind tunnel test, acquisition process test data；
2k. dismantles experimental rig according to the sequence opposite with installation, until the aircraft precursor 6 after the correction of the flank shape of 0.0D is disassembled,
It prepares for the aircraft precursor 6 after replacing the correction of the flank shape of 0.8D, 1.6D and 2.2D, wherein six component ring type balances 4 and gas supply turn
Connection between extension bar 5 does not have to disassembly, to reduce workload；
When 2l. replaces the aircraft precursor 6 after the correction of the flank shape of 0.8D, 1.6D and 2.2D, step 2c~2i is repeated, is then carried out corresponding
Wind tunnel test.
The present invention is not limited to above-mentioned specific embodiment, person of ordinary skill in the field from the above idea,
Without creative labor, made various transformation are within the scope of the present invention.
Claims (4)
1. being used for the flow tunnel testing device of air suction type hypersonic vehicle radome fairing separation simulation, it is characterised in that: the test
Device includes model equipment and jet apparatus；
The model equipment includes radome fairing head (1), radome fairing tail portion (2), through-hole cover board (3), six component ring type balances
(4), aircraft precursor (6), the aircraft fuselage (7) after correction of the flank shape, inner flow passage baffle after supplying adapting rod (5), correction of the flank shape
(8), support sting (9) and balance clamp nut (10)；The inner flow passage baffle (8) and support sting (9) is integrally machined into whole
Body, the aircraft fuselage (7) after correction of the flank shape are mounted on inner flow passage baffle (8), and the aircraft precursor (6) after correction of the flank shape, which is mounted on, to be repaired
On aircraft fuselage (7) after shape, gas supply adapting rod (5) is mounted on the aircraft precursor after correction of the flank shape (6), six component ring type days
Flat (4) are mounted on gas supply adapting rod (5), and radome fairing tail portion (2) are mounted on six component ring type balances (4), radome fairing head
(1) it is mounted on radome fairing tail portion (2), through-hole cover board (3) is mounted on radome fairing head, constitutes model equipment；
The jet apparatus includes aircraft precursor (6), the copper tube after retro-nozzle (11), gas supply adapting rod (5), correction of the flank shape
Road (12) and the reversed nut of screw thread (13)；Red copper pipeline (12) is mounted on the rear end of the aircraft precursor (6) after correction of the flank shape, and gas supply turns
Extension bar (5) is mounted on the front end of the aircraft precursor (6) after correction of the flank shape, and retro-nozzle (11) is installed by the reversed nut of screw thread (13)
On gas supply adapting rod (5), each contact surface is all made of copper washer sealing, constitutes jet apparatus；Red copper pipeline (12) and outer
The connection of boundary's gas source, imported into the aircraft precursor (6) after correction of the flank shape for the normal temperature compressed air of extraneous gas source, through supplying adapting rod
(5) enter retro-nozzle (11).
2. the flow tunnel testing device according to claim 1 for air suction type hypersonic vehicle radome fairing separation simulation,
It is characterized in that: having two groups of mounting-positioning holes on the ring flange of described gas supply adapting rod (5) rear end, the uniformly distributed interior hexagonal of inner ring is heavy
Hole and pin hole are used to for six component ring type balances (4) being fixed on gas supply adapting rod (5), the uniformly distributed interior hexagonal counterbore in outer ring and
Pin hole will be for that will supply the aircraft precursor (6) that adapting rod (5) is fixed on after correction of the flank shape.
3. the flow tunnel testing device according to claim 1 for air suction type hypersonic vehicle radome fairing separation simulation,
Be characterized in that: the aircraft precursor (6) after the correction of the flank shape has 4, the separation that the aircraft precursor (6) after each correction of the flank shape is adjusted
Distance respectively 0.0D, 0.8D, 1.6D and 2.2D, D is aircraft fuselage diameter, and the aircraft precursor (6) after each correction of the flank shape is internal
Have for vent hole, be connected to for one end of vent hole with red copper pipeline (12), is connected for other end gas supply adapting rod (5) of vent hole
It is logical.
4. the flow tunnel testing device according to claim 1 for air suction type hypersonic vehicle radome fairing separation simulation,
It is characterized in that: the conical nozzle beveling of the retro-nozzle (11), the angle and the cone angle phase of radome fairing head (1) of beveling
Together；The short side exit cross-section of retro-nozzle (11) and the relationship of nozzle throat cross section meet that jet flow pressure ratio is equal and jet flow is dynamic
Measure similar simulation rules.
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| CN201811119225.3A CN109250149B (en) | 2018-09-26 | 2018-09-26 | Wind tunnel test device for separation simulation of air suction type hypersonic vehicle fairing |
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| CN114993600A (en) * | 2022-07-18 | 2022-09-02 | 中国航天空气动力技术研究院 | Air-breathing type aircraft model fairing rapid separation device and separation method |
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| CN113588199A (en) * | 2021-09-30 | 2021-11-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Reverse jet flow blockage degree test device and test method for hypersonic wind tunnel |
| CN113588200B (en) * | 2021-09-30 | 2021-12-07 | 中国空气动力研究与发展中心超高速空气动力研究所 | High-flow reverse jet test device and method for hypersonic aircraft |
| CN113588200A (en) * | 2021-09-30 | 2021-11-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | High-flow reverse jet test device and method for hypersonic aircraft |
| CN114018527A (en) * | 2021-11-09 | 2022-02-08 | 中国空气动力研究与发展中心超高速空气动力研究所 | Design method of semi-automatic interactive wind tunnel test scheme |
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| CN114993600A (en) * | 2022-07-18 | 2022-09-02 | 中国航天空气动力技术研究院 | Air-breathing type aircraft model fairing rapid separation device and separation method |
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