A high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic heating constraint

Abstract

The application discloses a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic heat constraint, determines a rudder leading edge radius and a rudder leading edge sweepback angle based on a current state nominal trajectory and a rudder leading edge thermal response temperature as a judgment standard; secondly, a conical angle of a folding rudder-cone surface is determined through iterative calculation, so that a chordwise dimension of a high-temperature-resistant refractory metal material, and a thickness of a skin and a heatproof coating at the cone are designed; thirdly, through interference position design of a heatproof cup bow shock wave, a gap between a fixed rudder and a movable rudder avoids the bow shock wave interference zone, so that a chordwise distance between a folding gap and a rudder shaft is determined. Through the aerodynamic configuration design method, the rudder leading edge radius can be as low as 3mm, and the cone angle can be as low as 4°, so that ablation deformation does not occur when the maximum heat flow of the rudder leading edge is less than 10MW / m 2 2, the resistance characteristics of the high-speed aircraft can be effectively reduced, the aircraft has a near-zero ablation feature, and a good aerodynamic shape is maintained.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aircraft, and particularly relates to a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic heat constraint. BACKGROUND

[0002] At present, the development trend of various aircrafts is to have a longer range, a smaller size and a lighter mass. In order to facilitate transportation and storage, the aircraft and its launching device must be miniaturized, and therefore the aircraft needs to adopt a foldable rudder wing.

[0003] The folding rudder, as a key component of flight control, bears extremely severe aerodynamic heat and aerodynamic force load conditions in the process of high-speed flight. The conventional high-speed aircraft folding rudder adopts an ablative heat protection scheme. Due to the limitation of the production process of the composite material, the folding rudder usually has a large leading edge radius, which is a compromise of drag reduction target and heat protection. The folding rudder cannot adapt to the development trend of high-speed aircrafts. In addition, the leading edge of the ablative folding rudder will be severely ablated, which objectively causes the continuous increase of the folding rudder resistance. Therefore, it is a key problem to be solved to research a low-resistance non-ablation folding rudder suitable for high-speed aircrafts. The design process of the non-ablation folding rudder usually guarantees to meet the structural space requirements, and then meets the basic aerodynamic performance indicators. The design process is usually that the aerodynamic department proposes an initial shape, and then the aerodynamic heat environment is evaluated. If the aerodynamic heat response temperature does not meet the use requirements of the structural material, the folding rudder aerodynamic shape needs to be improved. After several iterations of the design cycle, the final folding rudder shape is obtained.

[0004] This design method mainly based on aerodynamic force and supplemented by aerodynamic heat is not optimal in design results and design cycle. The main shortcomings are as follows: first, the design of the tail rudder shape is mainly based on meeting the requirements of the aerodynamic performance indicators. The result of the aerodynamic design is often only to meet the basic aerodynamic performance indicators, and the structural heat response temperature is not relatively optimal. Second, the design cycle is long. Usually, the structural design personnel only design the tail rudder shape by meeting the aerodynamic performance indicators, and cannot meet the allowable requirements of the use material temperature. This will lead to the lengthening of the design iteration cycle of aerodynamic force and aerodynamic heat. Therefore, how to simplify the design process, shorten the design cycle, and simultaneously consider the aerodynamic heat and aerodynamic force performance of the folding rudder has become a technical problem to be solved in the aerodynamic design of the supersonic non-ablation transverse folding tail rudder. SUMMARY

[0005] In order to overcome the prior art, the present application provides a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic heat constraint, which is based on the current state nominal trajectory, and the rudder leading edge heat response temperature is used as the judgment standard to determine the rudder leading edge radius and the rudder leading edge sweepback angle; secondly, the cone angle of the folding rudder-cone surface is determined through iterative calculation, so as to design the chordwise size of the high-temperature refractory metal material, and the thickness of the skin and the heat-proof coating at the cone; thirdly, the interference position of the heat-proof cup bow shock is designed, so that the gap between the fixed rudder and the movable rudder avoids the bow shock interference area, thereby determining the chordwise distance between the folding gap and the rudder shaft. Through the aerodynamic configuration design method, the rudder leading edge radius can be as low as 3mm, and the cone angle can be as low as 4°, which can ensure that the maximum heat flow of the rudder leading edge does not exceed 10MW / m 2 When the ablation does not occur, the resistance characteristics of the high-speed aircraft can be effectively reduced. Compared with the traditional heat-proof sleeve + skin skeleton high-speed air rudder, the present application has the characteristics of near-zero ablation and maintains good aerodynamic shape.

[0006] The technical solution adopted by the present application to solve the technical problems is as follows:

[0007] Step 1: initial rudder surface shape design;

[0008] According to the overall index, the tail rudder index is decomposed and designed, and the decomposition index is: full-missile lift-drag characteristics, minimum static stability characteristics, maximum lift-drag ratio, control efficiency and maximum overload; according to the folding size limit, the maximum span of the rudder surface is determined, and then according to the minimum static stability characteristics, the projection area of the rudder surface is determined, and according to the initial leading edge sweepback angle of the rudder surface, the rudder surface root chord length and tip chord length are determined;

[0009] Step 2: rudder surface leading edge radius and sweepback angle design;

[0010] The rudder leading edge adopts high-temperature refractory metal material, and a design allowance is reserved. According to the heat response temperature at the stagnation point, the rudder surface leading edge fillet radius and the leading edge sweepback angle are inversely designed;

[0011] Step 3: rudder surface cone angle design;

[0012] The highest temperature on the rudder surface cone is calculated through iterative calculation, and the bevel angle at the rudder surface cone is inversely calculated, so as to design the chordwise length of the inlay strip, and the thickness of the skin and the heat-proof coating at the cone;

[0013] Step 4: fixed rudder and movable rudder joint gap design;

[0014] Through the simulation analysis of the heat-proof cup of the rudder shaft, the interference area of the bow shock generated by the heat-proof cup is calculated, so that the gap between the fixed rudder and the movable rudder avoids the bow shock, thereby determining the distance between the fixed rudder gap and the heat-proof cup of the rudder shaft, and the chordwise length of the fixed rudder is obtained.

[0015] Preferably, the step 2 is specifically:

[0016] According to the wide speed domain and wide air domain flight envelope of high-speed aircraft, the engineering calculation formula of the control surface stagnation point heat flux density is introduced, as shown below:

[0017]

[0018] Wherein: q represents the heat flux density on the leading edge stagnation line; R represents the control surface leading edge stagnation point radius; p ∞ represents the incoming flow atmospheric density, v ∞ represents the incoming flow velocity, h w represents the wall enthalpy, h s represents the stagnation enthalpy, n represents the power coefficient, ∧ e represents the effective sweepback angle of the control surface leading edge;

[0019] Taking the material allowable temperature as the limit temperature, the maximum heat flux at the control surface leading edge is calculated reversely, so as to optimize and iteratively select the reverse radius of the control surface leading edge and the leading edge sweepback angle.

[0020] Preferably, the step 3 is specifically:

[0021] The wing surface heat flux calculation adopts the strip method, which divides the wing surface into many strips parallel to the incoming flow direction, and assumes that each spanwise section has no influence on each other, so as to convert the three-dimensional problem into a two-dimensional problem for calculation;

[0022] The laminar and turbulent Eckert flat plate reference enthalpy method is adopted, and the laminar heat flux calculation formula is:

[0023]

[0024] In the formula, q w,l represents the laminar heat flux density, P r represents the Prandtl number, p e represents the density of the boundary layer outer edge, u e represents the velocity of the boundary layer outer edge, Re s represents the stagnation point Reynolds number, p * represents the density at the reference temperature, μ * represents the viscosity coefficient at the reference temperature, μ e represents the viscosity coefficient of the boundary layer outer edge;

[0025] The turbulent heat flux calculation formula is:

[0026]

[0027] In the formula, q w,t represents the turbulent heat flux density, C c represents the compression factor;

[0028] The oblique shock wave interference area of the rudder surface is calculated according to the blunt leading edge flat plate, and the expression of the oblique shock wave and a conical deflection angle is as follows

[0029]

[0030] In the formula, M1 is the Mach number before the shock wave, beta is the oblique shock angle, theta is the flow deflection angle, and gamma is the specific heat ratio.

[0031] Preferably, the step 4 is specifically as follows:

[0032] The interference area of the bow shock wave generated by the heat shield cup is calculated, so that the gap between the fixed rudder and the movable rudder avoids the bow shock wave, and the distance between the bow shock wave and the heat shield cup is expressed as follows:

[0033]

[0034] Wherein: D represents the diameter of the heat shield cup, which needs to be determined according to the heat protection material system of the heat protection structure; Ma represents the free stream Mach number, and the maximum position of the bow shock wave away from the rudder shaft is calculated according to different Mach numbers, so as to avoid the interference of the shock wave on the gap between the movable rudder and the fixed rudder.

[0035] Preferably, the effective sweep angle of the leading edge satisfies 30≤∧ e ≤60, and n=1.5.

[0036] Preferably, the high-temperature refractory metal material is molybdenum alloy or niobium tungsten alloy.

[0037] The beneficial effects of the present application are as follows:

[0038] The present application provides a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic heat constraint, designs a non-ablation rudder with excellent aerodynamic performance and light structure, and has wide application prospect.

[0039] (1) The material of the rudder leading edge is high-temperature refractory metal material, the rudder leading edge radius can be as low as 3mm through comprehensive optimization design of the rudder leading edge radius and sweep angle, the leading edge sweep angle is not greater than 55°, and the ablation deformation of the rudder leading edge can be prevented when the maximum heat flow of the rudder leading edge is not more than 10MW / m 2 ;

[0040] (2) The overall thickness of the folding rudder is small, the rudder surface conical bevel angle is not greater than 4°, and the ablation deformation does not occur, and the aerodynamic resistance is small.

[0041] (3) The gap between the fixed rudder and the movable rudder effectively avoids the interference area of the bow shock wave generated by the heat shield cup of the rudder surface, and the interference of the shock wave on the gap between the fixed rudder and the movable rudder is reduced. DETAILED DESCRIPTION

[0042] Figure 1The method flow chart of the present application.

[0043] Figure 2 The schematic diagram of the unfolded state profile of the folding rudder.

[0044] Reference signs: 1, rudder surface front edge rounding radius; 2, rudder surface front edge sweepback angle; 3, rudder surface front edge inlay strip; 4, rudder surface one taper (including skin and coating thickness); 5, gap between fixed rudder and movable rudder. DETAILED DESCRIPTION

[0045] The present application is further illustrated below in combination with the drawings and examples.

[0046] As shown in Figure 1 and Figure 2 , a high-speed aircraft folding rudder aerodynamic configuration design method based on aerodynamic heat constraint, the specific steps are as follows:

[0047] Step one: initial rudder surface profile design;

[0048] According to the overall indicators, the tail rudder indicators are decomposed and designed, and the decomposition indicators are: aircraft lift-drag characteristics, minimum static stability characteristics, maximum lift-drag ratio, control efficiency and maximum overload. The key is to determine the maximum span of the rudder surface according to the folding size limit, and then to determine the projected area of the rudder surface according to the minimum static stability characteristics, and to determine the rudder surface root chord length and tip chord length according to the initial front edge sweepback angle of the rudder surface.

[0049] Step two: rudder surface front edge radius and sweepback angle design;

[0050] The rudder front edge adopts high-temperature-resistant refractory metal material, and a design allowance is reserved. According to the thermal response temperature at the stagnation point, the rudder surface front edge rounding radius and the front edge sweepback angle are inversely designed.

[0051] Step three: rudder surface one taper angle design;

[0052] The highest temperature on the rudder surface one taper is calculated through iteration, the bevel angle at the rudder surface one taper is inversely calculated, and thus the chordwise length of the inlay strip and the thickness of the skin and heat-resistant coating at the one taper are designed.

[0053] Step four: fixed rudder and movable rudder joint gap design;

[0054] The radius size of the heat-resistant cup of the rudder shaft is simulated, the interference area of the bow shock wave generated by the heat-resistant cup is calculated, the gap between the fixed rudder and the movable rudder is avoided from the bow shock wave, and thus the distance between the fixed rudder gap and the heat-resistant cup of the rudder shaft is determined, and the chordwise length of the fixed rudder is obtained.

[0055] Embodiment:

[0056] Taking a typical wingless high-speed aircraft layout form as an example, the present application is further described in detail:

[0057] Step one, initial rudder design;

[0058] According to the overall indicators, the tail rudder index is decomposed and designed, and the decomposition indicators are: total missile lift-drag characteristics, minimum static stability characteristics, maximum lift-drag ratio, control efficiency and maximum overload. The key is to determine the maximum span of the rudder according to the folding size limit, and then to determine the projected area of the rudder according to the minimum static stability characteristics, and to determine the rudder root chord length and tip chord length according to the initial rudder leading edge sweep angle.

[0059] Step two, rudder leading edge radius and sweep angle design;

[0060] According to the wide speed range and wide air range flight envelope of the high-speed aircraft, in order to quickly iterate the rudder aerodynamic heat, the engineering calculation formula of the missile body stagnation point heat flux density function is introduced, as follows:

[0061]

[0062] In order to ensure that the rudder leading edge still has sufficient strength and shape under the action of ultra-high temperature thermal load, the material of the rudder leading edge is high-temperature refractory metal, which can be selected from molybdenum alloy, niobium tungsten alloy, etc. The maximum heat flux at the rudder leading edge is calculated inversely from the allowable temperature of the material as the limit temperature, so as to optimize and select the rudder leading edge reverse radius and leading edge sweep angle.

[0063] Step three, rudder one cone angle design;

[0064] In order to verify the chordwise length of the rudder leading edge strip, the aerodynamic heat on the rudder one cone surface needs to be iteratively calculated to meet the heat flux density constraint condition, and the rudder one cone angle is determined.

[0065] The wing surface heat flux can be calculated by the strip method. The wing surface is divided into many strips parallel to the flow direction, and it is assumed that each spanwise section has no effect on each other, and the three-dimensional problem is converted into a two-dimensional problem for calculation. From the overall point of view, the windward surface of the wing is quite flat, and obviously it is appropriate to use blunt leading edge flat heat flow formula along the strip, and here the Eckert flat reference enthalpy method for laminar and turbulent flow is selected. The calculation formula for laminar heat flow is:

[0066]

[0067] The calculation formula for turbulent heat flow is:

[0068]

[0069] The rudder oblique shock interference area is calculated according to the blunt leading edge flat plate, mainly according to the flow deflection angle of the opposite one cone, and the expression of the rudder oblique shock and the one cone deflection angle is as follows:

[0070]

[0071] Step four, rudder shaft anti-heat cup interference area;

[0072] The principle of designing the gap between the fixed rudder and the movable rudder is that the gap between the fixed rudder and the movable rudder is kept away from the interference area of the bow shock wave generated by the heat shield cup by calculating the interference area of the bow shock wave generated by the heat shield cup. The distance between the bow shock wave and the heat shield cup is expressed as follows:

[0073]

[0074] Wherein: D represents the diameter of the heat shield cup, which needs to be determined according to the heat protection material system of the heat shield structure; Ma represents the free stream Mach number, and the maximum position of the bow shock wave away from the rudder shaft is calculated according to different Mach numbers, so as to avoid the interference of the shock wave on the gap between the movable rudder and the fixed rudder.

Claims

1. A method for designing the aerodynamic configuration of a high-speed aircraft folding rudder based on aerodynamic thermal constraints, characterized in that: The steps include: Step 1: Initial rudder surface shape design; The tail rudder index is decomposed and designed based on the overall index, which includes: lift-drag characteristics of the whole missile, minimum static stability characteristics, maximum lift-drag ratio, control efficiency and maximum overload; Based on the folded size limit, the maximum span of the rudder surface is determined. Then, based on the minimum static stability characteristics, the projected area of ​​the rudder surface is determined. Based on the initial leading edge sweep angle of the rudder surface, the root chord length and tip chord length of the rudder surface are determined. Step 2: Design the rudder's front edge radius and sweep angle; The leading edge of the rudder is made of high-temperature resistant refractory metal material, with a design margin reserved. The radius of the rudder's leading edge fillet and the leading edge sweep angle are inversely designed based on the thermal response temperature at the stagnation point. Step 3: Design of the rudder surface cone angle; By iteratively calculating the maximum temperature on the first cone of the rudder surface, the bevel angle at the first cone of the rudder surface is calculated inversely, thereby designing the chord length of the slats, as well as the thickness of the skin and thermal protection coating at the first cone. Step 4: Design the gap between the fixed rudder and the dynamic rudder; Through simulation analysis of the rudder shaft heat shield, the interference area of ​​the bow shock wave generated by the heat shield is calculated, so that the gap between the fixed rudder and the dynamic rudder avoids the bow shock wave. The distance between the fixed rudder gap and the rudder shaft heat shield is determined, and the chord-wise length of the fixed rudder is obtained.

2. The aerodynamic configuration design method of a high-speed aircraft folding rudder based on aerodynamic thermal constraints according to claim 1, characterized in that: The step 2 is specifically as follows: According to the wide-speed and wide-airspace flight envelope of high-speed aircraft, the engineering calculation formula for the heat flux density at the stagnation point of the control surface is introduced as follows: Where: q represents the heat flux density on the leading edge stagnation point line; R represents the stagnation point radius of the rudder front edge; ρ ∞ represents the incoming atmospheric density, v ∞ Indicates the incoming flow velocity, h w represents the wall enthalpy, h s represents the stagnation enthalpy value, n represents the power coefficient, ∧ e Indicates the effective sweep angle of the rudder leading edge; Taking the allowable temperature of the material as the limit temperature, the maximum heat flux at the leading edge of the rudder is calculated, and the fillet radius and leading edge sweep angle of the rudder leading edge are optimized and iteratively selected.

3. The aerodynamic configuration design method of a high-speed aircraft folding rudder based on aerodynamic thermal constraints according to claim 2, characterized in that: The step 3 is specifically as follows: The strip method is used to calculate the heat flux on the airfoil surface. The airfoil surface is divided into many strips parallel to the incoming flow direction. It is assumed that the spanwise sections have no influence on each other, and the three-dimensional problem is converted into a two-dimensional problem for calculation. Using the laminar and turbulent Eckert plate reference enthalpy method, the laminar heat flow calculation formula is: Where q w,l represents the laminar heat flux, P r represents the Prandtl number, ρ e represents the density at the outer edge of the boundary layer, u e Re represents the velocity at the outer edge of the boundary layer, s is the stagnation point Reynolds number, ρ * represents the density at the reference temperature, μ * Indicates the viscosity coefficient at the reference temperature, μ e represents the viscosity coefficient at the outer edge of the boundary layer; The calculation formula for turbulent heat flow is: Where q w,t represents the turbulent heat flux, C c represents the compression factor; The oblique shock wave interference area of ​​the rudder surface is calculated based on the blunt leading edge flat plate. The expression of the oblique shock wave of the rudder surface and the deflection angle of a cone is as follows Where M1 is the Mach number of the shock front, β is the oblique shock angle, θ is the flow deflection angle, and γ is the specific heat ratio.

4. The aerodynamic configuration design method of a high-speed aircraft folding rudder based on aerodynamic thermal constraints according to claim 3, characterized in that: The step 4 is specifically as follows: By calculating the interference area of ​​the bow shock wave generated by the heat shield, the gap between the fixed rudder and the dynamic rudder can be made to avoid the bow shock wave. The distance between the bow shock wave and the heat shield is expressed as follows: Where: D represents the diameter of the heat protection cup, which needs to be determined according to the thermal protection material system of the heat protection structure; Ma represents the free flow Mach number. According to different Mach numbers, the maximum position of the bow shock wave from the rudder axis is calculated to avoid the shock wave interfering with the gap between the moving rudder and the static rudder.

5. The aerodynamic configuration design method of a high-speed aircraft folding rudder based on aerodynamic thermal constraints according to claim 4, characterized in that: The effective sweep angle of the leading edge satisfies 30≤∧ e When ≤60, n=1.

5.

6. The method for designing aerodynamic configuration of a folding rudder of a high-speed aircraft based on aerodynamic thermal constraints according to claim 5, characterized in that: The high temperature resistant refractory metal material is a molybdenum alloy or a niobium tungsten alloy.

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