CN106528990B - A kind of hypersonic pointed cone shape heat flow density modeling method based on functional optimization - Google Patents
A kind of hypersonic pointed cone shape heat flow density modeling method based on functional optimization Download PDFInfo
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Abstract
The present invention relates to a kind of hypersonic pointed cone shape heat flow density modeling method based on functional optimization, include the following steps: using hypersonic calorimetric wind-tunnel, ground calorimetric test is carried out to the pointed cone model of n different contracting ratios of aircraft；Heat flux distribution rule of the n different scale models in hypersonic calorimetric wind-tunnel is obtained respectively and obtains heat flow density test value Qwi, wherein 1≤i≤n；The wind tunnel test parameter of hypersonic calorimetric wind-tunnel is adjusted, obtaining first group of heat flow density test value Qwij, j is wind tunnel test number；Obtain the heat flux distribution rule of first group of aircraft；The heat flux distribution rule Qwk of k group aircraft is obtained altogether；Using functional optimization algorithm, wind-tunnel qualitative variables a and mark mould parameter b are introduced, calculating is iterated to Qwk, seeks optimal spatial alternation, obtains pointed cone heat flow density model Qw.The invention avoids the one-sidedness of prior art modeling method, reduce the interference of human experience factor.
Description
Technical field
The present invention relates to a kind of hypersonic pointed cone shape heat flow density modeling methods based on functional optimization, belong to modeling
Field.
Background technique
High-altitude, it is hypersonic under the conditions of, aircraft surface can generate the Aerodynamic Heating with very strong destructive power, heat flow density
It is influenced, is predicted by many factors such as flight Mach number, local Reynolds number, the vibrational excitation of gas molecule, dissociation even ionization
Modeling is difficult.Pointed cone has basis for the design of entire aircraft as a kind of representative configuration, its heat flow density of accurate description
Guiding significance.
Currently, for hypersonic heat flow density, both at home and abroad all in engineering approximation theoretical method, fluid calculation numerical simulation
Three aspects such as method, ground/flight test procedure are studied.The U.S. has developed a series of since 1970s
Aerodynamic Heating engineering approximation method software (AEROHEAT, LATCH, MINVER, UNLATCH2, UNLATCH3 etc.) and fluid calculation number
Value analogy method software (DPLR, FUN3D, GASP, LAURA, US3D etc.) (reference: hypersonic and high-temperature aerodynamics (the
2 editions), Anderson, aerospace industry publishing house, in September, 2003)；And carry out large number of ground test to try with flying demonstration verifying
It tests, has been built up the more mature predicting means of flight mission profile aerodynamic force thermal environment.Europe, Japanese (JAXA) etc. are also all
Has the certain calculating of hypersonic heat flow density and test capability.Domestic Aerospace Science and Technology Corporation, Chinese Academy of Sciences's mechanics study
Institute, Chinese aerodynamic investigation and centre of development, China Aerospace aerodynamic investigation institute, the National University of Defense technology etc. research units are all
The hypersonic heat flow density for having developed oneself calculates software, possesses hypersonic hot-fluid test capability and the test of continuous improvement
Technology (refers to " Chinese scientific strategy fluid dynamics ", Science Press, in September, 2014).
It makes a general survey of both at home and abroad, although reliable heat flow density data can be obtained under local condition, due to different wind-tunnel
It is different to the simulation context of flow parameter, thus heat flow density reflects different physical mechanisms, heat flux distribution rule disunity,
Aircraft thermal environment models mainly according to certain experience and it is assumed that providing engineering approximation and fitting, and still shortage is objective, has
There is the theoretical method of certain universality.
How a kind of hypersonic pointed cone shape heat flux distribution mathematical model of accurate response is provided, be this field urgently
The technical issues of solution.
Summary of the invention
It is an object of the invention to overcome the deficiencies of the prior art and provide a kind of hypersonic pointed cones based on functional optimization
Shape heat flow density modeling method, avoids the one-sidedness of modeling method, reduces the interference of human experience factor.
The object of the invention is achieved by following technical solution:
A kind of hypersonic pointed cone shape heat flow density modeling method based on functional optimization is provided, is included the following steps:
(1) hypersonic calorimetric wind-tunnel is utilized, calorimetric examination in ground is carried out to the pointed cone model of n different contracting ratios of aircraft
It tests；
(2) heat flux distribution rule of the n different scale models in hypersonic calorimetric wind-tunnel is obtained respectively to obtain
Heat flow density test value Qwi, wherein 1≤i≤n；
(3) the wind tunnel test parameter for adjusting hypersonic calorimetric wind-tunnel, obtains first group of heat flow density test value Qwij, j
For wind tunnel test number；
(4) the heat flux distribution rule Qw1=fij of first group of aircraft is obtained1(T0,P0,H0,Hw,p,ρ,u,M,
ReL, Rex), wherein T0 is wind-tunnel total temperature, P0 is wind-tunnel stagnation pressure, H0 is wind-tunnel total enthalpy, Hw is wind-tunnel wall surface enthalpy, and p is incoming flow pressure
Power, ρ be carry out current density, u is velocity vector, M is Mach number, ReL is unit Reynolds number, Rex is local Reynolds number；
(5) k-1 different hypersonic calorimetric wind-tunnel of replacement, repeat step (1)-(4), obtain k group aircraft altogether
Heat flux distribution rule Qwk=fijk(T0,P0,H0,Hw,p,ρ,u,M,ReL,Rex)；
(6) functional optimization algorithm is applied, wind-tunnel qualitative variables a and mark mould parameter b is introduced, calculating is iterated to Qwk, is asked
Optimal spatial alternation is taken, the expression formula of pointed cone heat flow density model Qw=f (t) is obtained.
Preferably, further include step (7), calculate according to Qw=f (t) the calculating data of acquisition, tried with practical flight
Data comparison is tested, when meeting required precision, determines Qw function expression；When precision is unsatisfactory for required precision, return step
(6), the function expression of Qw is regained.
Preferably, k different hypersonic calorimetric wind-tunnel include shock tunnel and gun wind tunnel.
Preferably, functional optimization algorithm is applied in step (6), introduces wind-tunnel qualitative variables a and mark mould parameter b, to Qwk into
Row iteration calculates, and obtains pointed cone heat flow density and calculates model Qw's method particularly includes: is calculated in function space using functional optimization
Method seeks f (t) the function representation form with least complex.
Preferably, f (t) the function representation shape with least complex is sought using functional optimization algorithm in function space
Formula method particularly includes: utilize the heat flux distribution rule Qwk=fij of k group aircraftk(T0,P0,H0,Hw,p,ρ,u,M,
ReL, Rex), substitute into formulaMost adaptation function shape is searched in canonical function library
Formula g (t)；
F (t) is preset solution functional form；G (t) is canonical function form；It is f (t), g (t) in Hilbert sky
Between angle；
G (t) and Qw has following relationship:
C+d/ (v+Rex)-Qw/ (ρ * u) (H0-Hw)=0
C, d, v are the constant in Qw function expression；Obtain the function expression of Qw are as follows: QW=St* (ρ * u) (H0-Hw),
St is because of Margoulis number.
The invention has the following advantages over the prior art:
(1) modeling method practicability of the invention is high, in modeling process, reflects the letter of unified heat flux distribution rule
Number is generated by functional optimization algorithm, and input parameter is only the high-precision differentiation Aerodynamic Heating for reflecting different physical mechanisms
Test data does not need other manual interventions.The one-sidedness for avoiding prior art modeling method, reduces human experience factor
Interference.
(2) there is better terseness using the model that the present invention obtains.Functional algorithm of the invention is in optimization process
Overall balance can be carried out between data residual sum model complexity, obtain succinct mathematical function analytical expression.
(3) there is the model obtained using the present invention better versatility ensure that by the test data of multiple wind-tunnel
The orthogonality of data ensure that the high-adaptability of model.The coverage that the present invention tests is high, to practical flight complexity Airflow Environment
Sufficiently simulation, ensure that the precision of prediction of heat flow density in actual flying test.
Detailed description of the invention
Fig. 1 is the modeling procedure figure of the modeling method of the invention；
Fig. 2 ground experiment point layout schematic diagram of the present invention；
Fig. 3 present invention models front and back heat flux distribution figure；Wherein Fig. 3 (a) is heat flux distribution figure before modeling；Fig. 3
It (b) is heat flux distribution figure after modeling.
Specific embodiment
Flow chart of the invention in different types of hypersonic wind tunnel as shown in Figure 1, carry out the point of different contracting ratios first
Calorimetric test is bored, obtains the high-precision differentiation aerothermodynamics experiment data for reflecting different physical mechanisms, and draw different tests
Heat flux distribution rule, judges whether rule is unified, if disunity, carries out spatial alternation under the guidance of functional optimization,
Continuous iteration, until obtaining the unified regularity of distribution.
Specific implementation step of the invention is provided below with reference to specific example:
(1) hypersonic calorimetric wind-tunnel is utilized, ground calorimetric is carried out to the pointed cone model of two different contracting ratios of aircraft
Test.Two kinds of contractings are than pointed cone model as shown in Fig. 2, carry out calorimetric test in wind-tunnel 1,2, and wherein test model is that semi-cone angle is
7 degree of pointed cone, 1 length L=571mm of model arrange 17 measuring points, and 2 length L=1136mm of model arranges 37 measuring points.
Obtain heat flow density data of the different scale models under different wind-tunnel difference operating conditions.Wherein, the trystate of wind-tunnel 1 is H0=
16MJ/kg, u=4979m/s, Thermochemical Non-equilibrium state；The trystate of wind-tunnel 2 be H0=3.3MJ/kg, u=2343m/s,
Calorimetric ideal gas.
(2) heat flux distribution rule of two different scale models in hypersonic calorimetric wind-tunnel is obtained respectively to obtain
Heat flow density test value Qw1 and Qw2；
(3) adjust j-1 hypersonic calorimetric wind-tunnel wind tunnel test parameter, including wind-tunnel total temperature T0, wind-tunnel stagnation pressure P0,
Wind-tunnel total enthalpy H0, wind-tunnel wall surface enthalpy Hw, incoming-flow pressure p, incoming flow density p, velocity vector u, Mach number M, unit reynolds number Re L,
Local reynolds number Re x obtains first group of heat flow density test value Qw1j and Qw2j；
(4) existing modeling method is utilized, the heat flux distribution rule Qw1=fij of first group of aircraft is obtained1(T0,
P0, H0, Hw, p, ρ, u, M, ReL, Rex), wherein T0 is wind-tunnel total temperature, P0 is wind-tunnel stagnation pressure, H0 is wind-tunnel total enthalpy, Hw is wind-tunnel
Wall surface enthalpy, p is incoming-flow pressure, ρ be carry out current density, u is velocity vector, M is Mach number, ReL be unit Reynolds number, Rex be work as
Ground Reynolds number；
(5) k-1 different hypersonic calorimetric wind-tunnel of replacement, repeat step (1)-(4), obtain k group aircraft altogether
Heat flux distribution rule Qwk=fijk(T0, P0, H0, Hw, p, ρ, u, M, ReL, Rex), type of wind tunnel should cover shock tunnel
And gun wind tunnel.
(6) functional optimization algorithm is applied, wind-tunnel qualitative variables a and mark mould parameter b is introduced, calculating is iterated to Qwk, is asked
Optimal spatial alternation is taken, is obtained pointed cone heat flow density model Qw=f (T0, P0, H0, Hw, p, ρ, u, M, ReL, Rex, a, b).
The function representation form with least complex is sought using functional optimization algorithm in function space；K group number
According to Qwk=fijk(T0, P0, H0, Hw, p, ρ, u, M, ReL, Rex) substitutes into formulaMost matching criteria functional form g (t) is searched in canonical function library, with
Value approach 0 most as optimal function form.
F (t) is preset solution functional form, f (t)=f (T0, P0, H0, Hw, p, ρ, u, M, ReL, Rex, a, b)；g
It (t) is canonical function form；For f (t), g (t) Hilbert space angle.
Most matching criteria functional form g (t) and Qw has following relationship:
C+d/ (v+Rex)-Qw/ (ρ * u) (H0-Hw)=0
C, d, v are the constant in Qw function expression.
Because of Margoulis number expression formula are as follows:
St=Qw/ (ρ * u) (H0-Hw)
Therefore St=c+d/ (v+Rex)
It acquires:
C=8.223e-4, d=220.1, v=3158.
Acquire QW=St* (ρ * u) (H0-Hw).
Shown in heat flux distribution rule such as Fig. 3 (b) after modeling, heat flux distribution figure such as Fig. 3 (a) institute before modeling
Show, the comparison of two width figures by a function this it appears that can only express the heat flux distribution rule of two wind-tunnel after modeling
Rule.
(7) the calculating data for calculate according to Qw=f (t) acquisition, it is smart when meeting with actual flying test data comparison
When degree requires, Qw function expression is determined；When precision is unsatisfactory for required precision, return step (6) regains the function of Qw
Expression formula.
The pointed cone shape heat flux distribution model that hypersonic wind tunnel modeling method of the invention obtains is applied to superb
In the Aerodynamic Heating high-precision forecast of velocity of sound flight, the heat flux distribution rule of Accurate Prediction aircraft outer surface different parts
Rule, precision of prediction is high, and complexity is low, and forecasting efficiency is high.
The above, optimal specific embodiment only of the invention, but scope of protection of the present invention is not limited thereto,
In the technical scope disclosed by the present invention, any changes or substitutions that can be easily thought of by anyone skilled in the art,
It should be covered by the protection scope of the present invention.
The content that description in the present invention is not described in detail belongs to the well-known technique of professional and technical personnel in the field.
Claims (3)
1. a kind of hypersonic pointed cone shape heat flow density modeling method based on functional optimization, it is characterised in that including walking as follows
It is rapid:
(1) hypersonic calorimetric wind-tunnel is utilized, ground calorimetric test is carried out to the pointed cone model of n different contracting ratios of aircraft；
(2) heat flux distribution rule of the n different scale models in hypersonic calorimetric wind-tunnel is obtained respectively obtains hot-fluid
Density test value Qwi, wherein 1≤i≤n；
(3) the wind tunnel test parameter of hypersonic calorimetric wind-tunnel is adjusted, obtaining first group of heat flow density test value Qwij, j is wind
Hole test number (TN)；
(4) the heat flux distribution rule Qw1=fij of first group of aircraft is obtained1(T0,P0,H0,Hw,p,ρ,u,M,ReL,
Rex), wherein T0 is wind-tunnel total temperature, P0 is wind-tunnel stagnation pressure, H0 is wind-tunnel total enthalpy, Hw is wind-tunnel wall surface enthalpy, and p is incoming-flow pressure, ρ
To carry out current density, u be velocity vector, M is Mach number, ReL is unit Reynolds number, Rex is local Reynolds number；
(5) k-1 different hypersonic calorimetric wind-tunnel of replacement, repeat step (1)-(4), obtain the hot-fluid of k group aircraft altogether
Density distributing law Qwk=fijk(T0,P0,H0,Hw,p,ρ,u,M,ReL,Rex)；
(6) functional optimization algorithm is applied, wind-tunnel qualitative variables a and mark mould parameter b is introduced, calculating is iterated to Qwk, is sought most
Excellent spatial alternation, acquisition pointed cone heat flow density model Qw=f (t)=f (T0, P0, H0, Hw, p, ρ, u, M, ReL, Rex, a,
b)；
It obtains pointed cone heat flow density and calculates model Qw's method particularly includes: seek to have using functional optimization algorithm in function space
There is f (t) the function representation form of least complex, method particularly includes: utilize the heat flux distribution rule Qwk of k group aircraft
=fijk(T0, P0, H0, Hw, p, ρ, u, M, ReL, Rex) substitutes into formulaIt is marking
Most adaptation function form g (t) is searched in quasi-function library；
F (t) is preset solution functional form；G (t) is canonical function form；It is f (t), g (t) in Hilbert space
Angle；
G (t) and Qw has following relationship:
C+d/ (v+Rex)-Qw/ (ρ * u) (H0-Hw)=0
C, d, v are the constant in Qw function expression；Obtain the function expression of Qw are as follows: QW=St* (ρ * u) (H0-Hw), St are
Because of Margoulis number.
2. the hypersonic pointed cone shape heat flow density modeling method as described in claim 1 based on functional optimization, feature
It is, further includes step (7), calculate according to Qw=f (t) the calculating data of acquisition, and actual flying test data comparison,
When meeting required precision, Qw function expression is determined；When precision is unsatisfactory for required precision, return step (6) is regained
The function expression of Qw.
3. the hypersonic pointed cone shape heat flow density modeling method as described in claim 1 based on functional optimization, feature
It is, k different hypersonic calorimetric wind-tunnel include shock tunnel and gun wind tunnel.
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| CN107808065B (en) * | 2017-11-23 | 2019-12-31 | 南京航空航天大学 | Three-dimensional complex-shape high-speed aircraft flow-solid-heat rapid calculation method |
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| CN112001034B (en) * | 2020-09-09 | 2022-03-29 | 中国空气动力研究与发展中心计算空气动力研究所 | Multi-surface cone-shaped flight wind tunnel model layout and design method |
| CN112378617B (en) * | 2020-12-09 | 2023-03-07 | 中国航天空气动力技术研究院 | Data processing method for eliminating interference of pulse wind tunnel heat measuring signal |
| CN114528645B (en) * | 2022-04-24 | 2022-07-01 | 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 hypersonic velocity aerodynamic thermal standard model for simulating three-dimensional complex flow |
| CN115312139B (en) * | 2022-09-23 | 2023-01-13 | 中国空气动力研究与发展中心计算空气动力研究所 | Method for accessing and converting hypersonic flow chemical reaction model data |
| CN117782515B (en) * | 2024-02-28 | 2024-05-07 | 中国空气动力研究与发展中心计算空气动力研究所 | Aerodynamic heat data uncertainty evaluation method for impact of shock tunnel inflow parameters |
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