CN111324931B - Interstage separation aerodynamic characteristic obtaining method for afterbody reverse jet flow - Google Patents
Interstage separation aerodynamic characteristic obtaining method for afterbody reverse jet flow Download PDFInfo
- Publication number
- CN111324931B CN111324931B CN202010163449.5A CN202010163449A CN111324931B CN 111324931 B CN111324931 B CN 111324931B CN 202010163449 A CN202010163449 A CN 202010163449A CN 111324931 B CN111324931 B CN 111324931B
- Authority
- CN
- China
- Prior art keywords
- flow
- jet
- inlet
- separation
- interstage separation
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Images
Classifications
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention relates to a method for acquiring interstage separation aerodynamic characteristics of rear body reverse jet flow, which comprises the steps of firstly fitting thrust characteristic data generated by ground test run of a jet flow engine into a curve as direct force and moment to be added into a rigid body dynamics equation, then solving the flow characteristics of a jet flow engine combustion chamber and a jet pipe at a typical moment, fitting flow parameters at the inlet of the jet pipe into the curve as the boundary condition of the inlet of the jet pipe when a computational fluid mechanics equation is solved, and finally coupling propulsion to solve the computational fluid mechanics-rigid body dynamics equation to obtain the interstage separation aerodynamic characteristics. The invention adopts two correction methods to solve the problem of prediction of the interstage separation aerodynamic characteristics under the condition of reverse jet interference, and the principle is that the inner flow and the outer flow are decomposed, the flow characteristics of a combustion chamber and a spray pipe are obtained in advance, then only the back-spray pipe is reserved, and the outer flow is coupled to solve.
Description
Technical Field
The invention relates to an interstage thermal separation aerodynamic characteristic acquisition method of an aircraft, which is used for simulating and solving interstage thermal separation aerodynamic design and aerodynamic characteristic evaluation of the aircraft with a backward jet flow.
Background
The interstage separation can be divided into cold separation and hot separation according to different separation power, wherein the cold separation refers to separation realized by pushing the two bodies away by means of pneumatic force or mechanical force, and the hot separation refers to a separation mode of separating the two bodies by means of high-temperature and high-pressure jet flow by means of front body or rear body engine jet flow.
The current numerical simulation technology, numerical simulation technology and wind tunnel test technology can better solve the problem of interstage cold separation of an aircraft, and the typical work includes that Gu Ruyan and other numerical simulation are performed on an axisymmetric elastomer interstage cold separation flow field under the condition of low altitude supersonic speed, so that the change law of front and rear body resistance characteristics is obtained (Gu Ruyan, jiang Zhenyu, hu Fan, zhang Weihua. Numerical simulation and resistance characteristics of axisymmetric elastomer interstage cold separation flow field [ J ], solid rocket technology, 2014,37 (5): 606-610); lv Yan et al studied the relative motion of the separation body under complex stress conditions using 6-degree-of-freedom simulation analysis software (Lv Yan, zhong Jiehua, yin Shiming. Interstage separation modeling and simulation based on uncertainty analysis [ J ], flight mechanics, 2019,37 (2): 72-76); song Wei et al proposed a multi-body separation wind tunnel test method for simulating engine residual thrust using a split latch mechanism and a spring housing as a driving device (inventor: song Wei, jiang Zenghui, gu Ouyao. Publication: cn104483088a. Patent name: a wind tunnel multi-body separation free flight test method for simulating engine residual thrust).
The interstage thermal separation of the aircraft puts forward a severe requirement on a ground test, firstly, high-temperature and high-pressure fuel gas cannot be directly provided by an air pump, the fuel gas cannot be stably conveyed into a wind tunnel according to a specified flow rate in a chemical reaction generation mode, and the erosion of the fuel gas flow to the wall surface of a spray pipe of the wind tunnel is very serious, so that the ground wind tunnel test of the thermal spraying separation is difficult to develop, and related invention patents cannot be inquired. The research work on thermal spray separation has focused on numerical simulation and numerical simulation, and is typified by: gu Ruyan and the like adopt a method for solving an axisymmetric unsteady Navier-Stroke equation and a one-dimensional separation kinetic equation in a coupling manner, and research interference flow field characteristics of thermal spray separation at the tail part of a low-altitude front body (Gu Ruyan, jiang Zhenyu and Zhang Weihua. Numerical simulation of flow field characteristics at the initial stage of low-altitude interstage thermal separation of a rocket [ J ], astronomical report, 2015,36 (11): 1310-1315); wu Yuyu A numerical simulation approach was used to analyze the pressure variation and transient pressure spike at the aft end frame during thermal spray separation of the precursor tail (Wu Yuyu, jiang Ping, high wave, any Peng. Interstage thermal separation transient high pressure spike research [ J ], missile and space vehicle technologies, 2017, 5. Through a great deal of research, the research of thermal spraying separation focuses on two-stage separation of a carrier rocket, the separation adopts high-temperature and high-pressure fuel gas sprayed from the tail of a front body to directly act on the head of a rear body to realize two-stage separation, and the head of the rear body needs to bear high thermal load and impact load.
The back body reverse spray type thermal separation has the advantages that the two bodies do not need to bear extremely high thermal load and impact load, the residual thrust of the back body engine can be effectively utilized, the fuel consumption of the front body engine is saved, and the application prospect is good. Compared with the traditional interstage separation solving method, the interstage separation with the reverse jet flow has two interference influences, on one hand, the reverse jet flow engine can generate extra thrust and moment on the rear body, and on the other hand, the reverse jet flow can generate new interference on the front and rear external flow fields. Therefore, the traditional method for solving (CFD-RBD) interstage separation by coupling the Navier-Stokes equation (RANS) and the 6-DOF rigid body motion equation (6 DOF-RBD) at unsteady Reynolds is not applicable any more, the influence of the internal flow of the combustion chamber of the engine needs to be considered, a chemical reaction equation of the combustion chamber needs to be added in the traditional solving method, the difficulty of the coupling solving of the three equations is very large, and the real appearance of the aircraft is very complex, so that the complete coupling solving method is unrealistic for the aerodynamic design and the aerodynamic characteristic prediction of the aircraft adopting the reverse jet interstage separation.
Disclosure of Invention
The invention solves the technical problems that: the method overcomes the defects of the prior art, provides a technology for acquiring the aerodynamic characteristics under the condition of thermal jet flow interstage separation, and meets the requirements of calculation and evaluation of the aerodynamic characteristics when the fuel gas reverse jet flow is used for aircraft interstage separation.
The technical solution of the invention is as follows: an interstage separation aerodynamic characteristic obtaining method aiming at afterbody reverse jet flow comprises the following steps:
1) Fitting thrust characteristic data generated by a ground trial run of a jet engine into a curve by adopting a least square method to obtain direct force F (t) and moment M (t) which change along with time;
2) Intercepting the shapes of a combustion chamber and a back-spray nozzle to calculate an individual internal flow field, selecting 3-5 typical moments from ground test run data of an engine according to the calculation state, and solving a steady computational fluid mechanics equation to obtain the pressure, the temperature and the speed at the inlet of the back-spray nozzle;
3) Fitting the pressure, temperature and speed at the inlet of the jet pipe of the jet flow engine into a curve by adopting a least square method to obtain the pressure P (T), the temperature T (T) and the speed V (T) of the inlet of the jet pipe, which change along with time;
4) Decomposing the direct force F (t) and the moment M (t) into a body axis system according to the attitude of the aircraft to obtain 6 components of Fx (t), fy (t), fz (t), mx (t), my (t) and Mz (t) under the body axis system; adding the 6 components as force and moment source terms into a rigid body dynamic equation;
5) Decomposing the inlet speed V (T) of the spray pipe into a calculation coordinate system according to the attitude of the aircraft to obtain Vx (T), vy (T) and Vz (T) under the calculation coordinate system, and adding inlet parameters P (T), T (T), vx (T), vy (T) and Vz (T) of the spray pipe into a computational fluid mechanics equation as inlet boundary conditions of the spray pipe;
6) And only the back-jet pipe and the external flow field are reserved, and the interstage separation hydrodynamics-rigid body dynamics coupling solving is carried out to obtain the separation motion rule and the pneumatic characteristic of the front body and the rear body.
And (2) in the step 1), the direct force F (t) and the moment M (t) are obtained, and curve fitting is carried out by adopting pneumatic characteristic data generated by a ground test of a jet engine.
And 3) in the step 3), curve fitting is carried out on the nozzle inlet pressure P (T), the temperature T (T) and the speed V (T) by adopting 3-5 typical moment steady flow field results.
In the step 4), the force and moment source terms Fx (t), fy (t), fz (t), mx (t), my (t) and Mz (t) required by solving the rigid body dynamic equation are obtained by adopting the fitting result in the step 1) to carry out axial system decomposition.
And (5) solving the required nozzle inlet boundary conditions P (T), T (T), vx (T), vy (T) and Vz (T) by using the fitting result in the step 3) to obtain the nozzle inlet boundary conditions through decomposition to a calculation coordinate system.
And 6) acquiring the inter-stage separation motion law and the pneumatic characteristic by adopting a CFD-RBD coupling solving method instead of an engineering estimation method or a numerical simulation method.
The interstage separation hydrodynamics-rigid body dynamics coupling solution in step 6) is to retain the flow inside the nozzle, rather than just the nozzle exit profile.
Compared with the prior art, the invention has the advantages that:
the invention provides an interstage separation aerodynamic characteristic obtaining method aiming at a afterbody reverse jet, which is characterized in that thrust borne by the afterbody is fitted into a curve as direct force and moment to be added into a rigid body dynamics motion equation, and flow parameters at the inlet of a jet pipe of a jet engine are fitted into a curve changing along with time to be used as a nozzle inlet boundary condition calculated by CFD (computational fluid dynamics-rigid body dynamics equation) during coupling solution of an interstage separation fluid dynamics-rigid body dynamics equation (CFD-RBD). Compared with the traditional thermal jet separation simulation method, the interstage separation aerodynamic characteristic obtaining method has the advantages that the influence of the internal flow of an engine combustion chamber does not need to be considered, an additional chemical reaction equation aiming at the combustion chamber does not need to be added, and the calculated amount is greatly reduced; the pneumatic interference generated by jet flow can be better simulated only by keeping the flow process in the spray pipe; little modification to the existing CFD-RBD solving program
Drawings
FIG. 1 is a flow chart of the present invention;
FIG. 2 (a) is an outline view of an interstage separation missile;
FIG. 2 (b) is a view of the combustion chamber and the back-nozzle profile;
FIG. 3 (a) shows the forward and aft centroids moving distance in the x direction of the incoming flow in 0-0.3 s;
FIG. 3 (b) shows the moving distance of the front and back body centroids in the y direction of the incoming flow in 0-0.3 s;
FIG. 3 (c) is a graph showing the distance the front and back body centroids move in the z direction of the incoming flow in 0-0.3 s;
FIG. 4 (a) shows the roll angle change of the anterior and posterior bodies between 0-0.3 s;
FIG. 4 (b) shows the change of yaw angle between 0s and 0.3s for the fore and aft bodies;
FIG. 4 (c) shows the change of pitch angle between 0s and 0.3s for the front and rear bodies.
Detailed Description
The flow chart of the invention is shown in figure 1 and is realized by the following steps:
1. according to thrust data (table 1) obtained by the ground test run of the reverse-injection engine, curve fitting is carried out to obtain direct force F (t) and moment M (t), and the fitting result is as follows:
F(t)＝25102e -8.417t (N)
M(t)＝0(N.m)
TABLE 1 ground test data for reverse-injection engines
| Serial number | Time t(s) | Pressure intensity P (MPa) of combustion chamber | Thrust of reverse-injection engine F (N) |
| 1 | 0 | 3.6 | 27000 |
| 2 | 0.03 | 2.9 | 19000 |
| 3 | 0.06 | 2.4 | 15000 |
| 4 | 0.09 | 1.8 | 11500 |
| 5 | 0.12 | 1.4 | 9000 |
| 6 | 0.15 | 1 | 7000 |
| 7 | 0.18 | 0.8 | 5500 |
| 8 | 0.21 | 0.6 | 4200 |
| 9 | 0.24 | 0.5 | 3300 |
| 10 | 0.27 | 0.4 | 2600 |
| 11 | 0.3 | 0.3 | 2100 |
2. The combustion chamber and the back-injection nozzle in FIG. 3 are intercepted for independent calculation, 3-5 typical moments are selected, and a steady RANS equation (computational aerodynamics, fu Dexun master edition, aerospace Press, 1994.11) is solved. In this example, 3 moments 0s, 0.15s and 0.3s in table 1 are selected, the corresponding combustion chamber pressures are 3.6MPa, 1MPa and 0.3MPa respectively, the physical parameters of the fuel gas are shown in table 2, and the flow parameters at the inlet of the back-spray nozzle are obtained, as shown in table 3:
TABLE 2 Combustion Chamber parameters
| Temperature (K) | Density (kg/m) 3 ) | Average molecular weight | Specific heat ratio |
| 3200 | 6 | 28 | 1.2 |
TABLE 3 nozzle inlet flow parameters
| Time t(s) | Pressure P (MPa) | Temperature T (K) | Speed V (m/s) |
| 0 | 1.47 | 3075 | 1293 |
| 0.15 | 0.44 | 3020 | 1250 |
| 0.3 | 0.13 | 2942 | 1216 |
3. The nozzle inlet parameters in table 3 were fitted to a time-varying curve with the following fit:
P(t)＝1e -0.8139t (MPa)，T(t)＝3078e -0.147t (K),V(t)＝1291e -0.205t (m/s)
4. taking the direct force F (t) and the moment M (t) obtained in the step 1 as source terms, adding the source terms into a rigid body motion equation with 6 degrees of freedom, and noting that in the propulsion solution of the rigid body dynamics equation, F (t) and M (t) are decomposed into a body axis system according to the attitude of an aircraft, namely, the actually adopted components are 6 in total, namely, fx (t), fy (t), fz (t), mx (t), my (t) and Mz (t) under the body axis system;
5. and (3) taking the nozzle inlet parameter pressure P (T), the temperature T (T) and the speed V (T) obtained in the step (3) as nozzle inlet boundary conditions for CFD calculation, and decomposing the nozzle inlet speed V (T) into a calculation coordinate system, namely actually adopting Vx (T), vy (T) and Vz (T) in the calculation coordinate system for CFD calculation.
6. Only the jet pipe and the external flow field are reserved, and interstage separation CFD-RBD coupling solving is carried out to obtain the separation motion law and the pneumatic characteristics of the front body and the rear body. The mass characteristics of the front and rear bodies are selected from table 4, the conditions of the incoming flow are calculated to be 30Km in height, 5.5 in mach number, 2 in attack angle, 1 in sideslip angle, and 0.3s in total time, fig. 3 shows the displacement of the mass center of the front and rear bodies relative to the initial position, and fig. 4 shows the attitude angle change of the front and rear bodies.
TABLE 4 Mass characteristics of front and rear bodies at interstage separation
The present invention has not been described in detail as is known to those skilled in the art.
Claims (7)
1. An interstage separation aerodynamic characteristic obtaining method aiming at afterbody reverse jet flow is characterized by comprising the following steps:
1) Fitting thrust characteristic data generated by ground test of a jet engine into a curve by adopting a least square method to obtain direct force F (t) and moment M (t) which change along with time;
2) Intercepting the shapes of a combustion chamber and a back-spray nozzle to calculate an individual internal flow field, selecting 3-5 typical moments from ground test run data of an engine according to the calculation state, and solving a steady computational fluid mechanics equation to obtain the pressure, the temperature and the speed at the inlet of the back-spray nozzle;
3) Fitting the pressure, temperature and speed at the inlet of the jet pipe of the jet flow engine into a curve by adopting a least square method to obtain the pressure P (T), the temperature T (T) and the speed V (T) of the inlet of the jet pipe, which change along with time;
4) Decomposing the direct force F (t) and the moment M (t) into a body axis system according to the attitude of the aircraft to obtain 6 components of Fx (t), fy (t), fz (t), mx (t), my (t) and Mz (t) under the body axis system; adding the 6 components as force and moment source terms into a rigid body dynamic equation;
5) Decomposing the inlet speed V (T) of the spray pipe into a calculation coordinate system according to the attitude of the aircraft to obtain Vx (T), vy (T) and Vz (T) under the calculation coordinate system, and adding inlet parameters P (T), T (T), vx (T), vy (T) and Vz (T) of the spray pipe into a computational fluid mechanics equation as inlet boundary conditions of the spray pipe;
6) Only the back-spraying pipe and the outer flow field are reserved, and inter-stage separation hydrodynamics-rigid body dynamics coupling solving is carried out to obtain the separation motion rule and the pneumatic characteristics of the front body and the rear body.
2. The method for obtaining the interstage separation aerodynamic characteristic of the afterbody reverse jet according to claim 1, characterized in that: and (2) in the step 1), the direct force F (t) and the moment M (t) are obtained, and curve fitting is carried out by adopting pneumatic characteristic data generated by a ground test of a jet engine.
3. The method for obtaining interstage separation aerodynamic characteristics for aft body reverse jets according to claim 1, wherein: and 3) in the step 3), the nozzle inlet pressure P (T), the temperature T (T) and the speed V (T) are obtained, and curve fitting is carried out by adopting the steady flow field results at 3-5 typical moments.
4. The method for obtaining interstage separation aerodynamic characteristics for aft body reverse jets according to claim 1, wherein: in the step 4), the force and moment source terms Fx (t), fy (t), fz (t), mx (t), my (t) and Mz (t) required by solving the rigid body dynamic equation are obtained by adopting the fitting result in the step 1) to carry out axial system decomposition.
5. The method for obtaining interstage separation aerodynamic characteristics for aft body reverse jets according to claim 1, wherein: and (5) solving the required nozzle inlet boundary conditions P (T), T (T), vx (T), vy (T) and Vz (T) by using the fitting result in the step 3) to obtain the nozzle inlet boundary conditions through decomposition to a calculation coordinate system.
6. The method for obtaining interstage separation aerodynamic characteristics for aft body reverse jets according to claim 1, wherein: and 6) acquiring the inter-stage separation motion law and the pneumatic characteristic by adopting a CFD-RBD coupling solving method instead of an engineering estimation method or a numerical simulation method.
7. The method for obtaining interstage separation aerodynamic characteristics for aft body reverse jets according to claim 1, wherein: the interstage separation hydrodynamics-rigid body dynamics coupling solution in step 6) preserves the flow inside the nozzle, rather than only the nozzle exit profile.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010163449.5A CN111324931B (en) | 2020-03-10 | 2020-03-10 | Interstage separation aerodynamic characteristic obtaining method for afterbody reverse jet flow |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010163449.5A CN111324931B (en) | 2020-03-10 | 2020-03-10 | Interstage separation aerodynamic characteristic obtaining method for afterbody reverse jet flow |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN111324931A CN111324931A (en) | 2020-06-23 |
| CN111324931B true CN111324931B (en) | 2022-10-28 |
Family
ID=71171455
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202010163449.5A Active CN111324931B (en) | 2020-03-10 | 2020-03-10 | Interstage separation aerodynamic characteristic obtaining method for afterbody reverse jet flow |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN111324931B (en) |
Families Citing this family (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112417775B (en) * | 2020-10-30 | 2024-03-26 | 中国运载火箭技术研究院 | Method for calculating ballistic performance in thrust termination and pressure relief process of solid rocket engine |
| CN112346359B (en) * | 2020-11-04 | 2022-05-31 | 中国空气动力研究与发展中心高速空气动力研究所 | Direct force-aerodynamic force composite control numerical simulation method based on CFD and RBD coupling |
| CN112528420B (en) * | 2020-12-25 | 2022-05-27 | 中国空气动力研究与发展中心计算空气动力研究所 | Dynamic boundary condition switching method for jet flow time sequence control simulation |
| CN114528775B (en) * | 2021-12-28 | 2024-06-25 | 中国航天空气动力技术研究院 | Numerical simulation method, device and equipment for rocket stage separation pressure-holding process |
| CN114992674B (en) * | 2022-06-20 | 2024-03-19 | 中国航发贵阳发动机设计研究所 | External relay ignition and starting oil supply device |
| CN117875222B (en) * | 2024-03-12 | 2024-05-17 | 中国空气动力研究与发展中心计算空气动力研究所 | Long-distance interstage separation track numerical value prediction method and equipment |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106019930A (en) * | 2016-08-03 | 2016-10-12 | 中国人民解放军63821部队 | Aerodynamic/control integrated coupling simulating technology in aircraft maneuvering process |
| CN109297666A (en) * | 2018-10-15 | 2019-02-01 | 中国空气动力研究与发展中心高速空气动力研究所 | A kind of stage separation flow tunnel testing device and test method based on two sets of movement mechanisms |
-
2020
- 2020-03-10 CN CN202010163449.5A patent/CN111324931B/en active Active
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106019930A (en) * | 2016-08-03 | 2016-10-12 | 中国人民解放军63821部队 | Aerodynamic/control integrated coupling simulating technology in aircraft maneuvering process |
| CN109297666A (en) * | 2018-10-15 | 2019-02-01 | 中国空气动力研究与发展中心高速空气动力研究所 | A kind of stage separation flow tunnel testing device and test method based on two sets of movement mechanisms |
Also Published As
| Publication number | Publication date |
|---|---|
| CN111324931A (en) | 2020-06-23 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN111324931B (en) | Interstage separation aerodynamic characteristic obtaining method for afterbody reverse jet flow | |
| Ding et al. | An overview of waverider design concept in airframe/inlet integration methodology for air-breathing hypersonic vehicles | |
| EP1746523B1 (en) | Method of designing a hypersonic inlet for an aircraft and system therefor | |
| Huang et al. | Investigation on high angle of attack characteristics of hypersonic space vehicle | |
| Naitoh et al. | A new cascade-less engine operated from subsonic to hypersonic conditions: designed by computational fluid dynamics of compressible turbulence with chemical reactions | |
| Gamble et al. | Improving off-design nozzle performance using fluidic injection | |
| Surber et al. | Inlet flow control technology: learning from history, reinventing the future | |
| Korzun et al. | Application of a detached eddy simulation approach with finite-rate chemistry to mars-relevant retropropulsion operating environments | |
| Li et al. | Numerical investigations on the aerodynamics of SHEFEX-III launcher | |
| Srivastava | Aerodynamic performance of supersonic missile body-and wing tip-mounted lateral jets | |
| Cockrell Jr | Interpretation of waverider performance data using computational fluid dynamics | |
| Fei et al. | Aerodynamic characteristics of store during lateral jet assisted separation from cavity using free drop technique | |
| Murty et al. | Numerical characterisation of jet-vane based thrust vector control systems | |
| Rodi | Preliminary ramjet/scramjet integration with vehicles using osculating flowfield waverider forebodies | |
| Dunsworth et al. | Ramjet engine testing and simulation techniques | |
| Lv et al. | Design and performance analysis of an exhaust system for an over-under turbine based combined cycle operating at Mach 0–6 | |
| Pandya et al. | Capsule Abort Recontact Simulation | |
| Singh et al. | Computational fluid dynamics in aerospace industry in India | |
| Sanaka et al. | Re-entry vehicle performance analysis under the control of lateral jet | |
| Riggins et al. | Methods for the design of energy efficient high speed aerospace vehicles | |
| Srivastava | CFD analysis and validation of lateral jet control of a missile | |
| Maxwell et al. | Shapeable inlet manifold for hypersonic scramjet | |
| Lin et al. | Novel approach for maneuvering reentry vehicle design | |
| Lohner et al. | On the drag efficiency of counterjets in low supersonic flow | |
| Kim et al. | Performance Characteristics of Hypersonic External Compression Inlet Using Isentropic Compression Surface |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |