A design method of a new pre-compressor wing for hypersonic air-breathing vehicles
By designing the wing pre-compression angle, characteristic area, shape, and cross-section of the hypersonic air-breathing aircraft, the problem of unbalanced internal and external flow performance was solved, achieving high-efficiency aerodynamic and maneuverability performance within the same flight angle of attack range.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Filing Date
- 2023-12-27
- Publication Date
- 2026-05-15
AI Technical Summary
Existing research on the integration of internal and external flow in hypersonic air-breathing aircraft focuses primarily on the simulation and optimization of local structures and flow fields, while insufficient simulation of the internal flow indices of the wing's aerodynamic shape, resulting in a decrease in lift-to-drag ratio and reduced maneuverability.
By designing the wing's pre-compression angle, characteristic area, shape, and cross-section, and combining this with a coupled analysis of the variation of internal and external flow performance with the angle of attack, the wing's aerodynamic shape is adjusted to balance internal and external flow performance within the same flight angle of attack range.
This achievement enables hypersonic air-breathing aircraft to achieve high levels of internal and external flow performance within the same flight angle of attack range, thereby improving the aerodynamic and maneuverability of the aircraft.
Smart Images
Figure CN117775270B_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of aerodynamic shape design technology for hypersonic air-breathing aircraft, specifically to a design method for a novel pre-compressed wing for hypersonic air-breathing aircraft. Background Technology
[0002] Hypersonic air-breathing vehicles are aircraft powered by air-breathing scramjet engines. They combine the aerodynamic performance of the aircraft with the thrust performance of the air-breathing engine. With their advantages of high speed, long range and superior performance, they have been continuously studied by researchers at home and abroad. However, as the flight speed increases, the drag experienced by the aircraft increases sharply, resulting in a decrease in lift-to-drag ratio and reduced maneuverability. Therefore, promoting the integrated design of internal and external flow is the key to realizing hypersonic air-breathing flight. At present, the research on the integrated design of internal and external flow of aircraft still focuses more on the local structure and flow field simulation optimization, while the optimization of wing aerodynamic shape through internal flow index simulation data is still relatively rare.
[0003] Therefore, a novel pre-compressed wing for hypersonic air-breathing aircraft that balances internal and external flow performance is proposed to address the above problems. Summary of the Invention
[0004] The purpose of this invention is to provide a novel pre-compressed wing for hypersonic air-breathing aircraft that balances internal and external flow performance, in order to address the problem mentioned in the background art that existing research on the integration of internal and external flow in aircraft still focuses more on local structure and flow field simulation optimization, while the optimization of wing aerodynamic shape through internal flow index simulation data is still relatively limited.
[0005] To achieve the above objectives, the present invention provides the following technical solution: a novel pre-compressed wing for a hypersonic air-breathing aircraft that balances internal and external flow performance. The wing pre-compression angle, characteristic area, shape, and cross-section are based on a coupled analysis of the variation law of internal and external flow performance with angle of attack. Through aerodynamic design of the wing pre-compression angle and configuration, the internal and external flow performance is adjusted to achieve a high level within the same flight angle of attack range.
[0006] Preferably, the wing precompression angle θ pre-c This includes the characteristics of internal flow performance as a function of angle of attack, a reference lift-to-drag ratio, and the wing pre-compression angle θ. pre-c The angular difference between the wing chord and the fuselage axis on the longitudinal reference XOY plane, such as Figure 2 As shown, if the optimal internal flow performance corresponds to an angle of attack of α1, and the optimal external flow performance corresponds to an angle of attack of α2, then That is, the wing chord line must have a θ relative to the central axis of the aircraft. pre-c The pre-compression angle.
[0007] Preferably, the wing characteristic area is represented as Scharacter, where Scharacter is the area obtained by projecting the wing shape onto the fuselage normal. Based on the high Mach number compression-expansion aerodynamic principle, the pressure distribution on the windward / leeward side of the wing is estimated, such as... Figure 3 As shown, the normal aerodynamic load of the pre-compressed airfoil is then obtained through the pressure difference on both sides. Then, based on the optimal outflow performance corresponding to the flight angle of attack α2, the normal load can be decomposed into the lift and drag increments caused by the wing, and the lift increment. Incremental resistance Then, based on the initial lift-to-drag ratio and target lift-to-drag ratio Calculate the characteristic area of the wing The pressure difference between the windward and leeward sides of the wing is (P wind -P lee The angle of attack corresponding to the optimal outflow performance is α2, which can be calculated from the aircraft's extreme flight conditions.
[0008] Preferably, the wing shape includes a near-body side curve, a far-body side curve, and a trailing edge curve.
[0009] Preferably, the wing cross-section includes a thickness distribution and a passivated leading edge cross-sectional shape.
[0010] Preferably, the internal flow performance varies with the angle of attack, such as... Figure 4 As shown, the internal flow performance characteristics of the same type of inlet at different angles of attack are first obtained through CFD numerical simulation. Then, a moving interval is formed by three angle values, such as starting with (0... ° ,1 ° ,2 ° ) is calculated for the interval, then move to (1 ° ,2 ° ,3 ° Calculate, and so on, the moving averages of the total pressure recovery coefficient σ, the throat pressure ratio π, and the inlet capture flow rate m, respectively. With moving standard deviation Subsequently, three intervals with higher moving mean and lower moving standard deviation were identified (α). a ,α b ,α c ) σ ,(α d ,α e ,α f ) π ,(α g ,α h ,α i ) m The intersection of the three intervals represents the range of values for the optimal internal flow performance angle of attack α1.
[0011] Preferably, the reference lift-to-drag ratio includes an initial lift-to-drag ratio and a target lift-to-drag ratio.
[0012] Preferably, the near-fuselage side curve is the intersection curve of the wing and fuselage, therefore its shape is consistent with the curve shape at the corresponding position on the fuselage, and can be fitted using a cubic curve, i.e., curve AB is... .
[0013] Preferably, the remote fuselage side curve AC can be a quadratic curve. Perform fitting.
[0014] Preferably, the trailing edge curve BC can be a quadratic curve. To perform fitting, establish a rectangular coordinate system with point A as the origin, such as... Figure 5 As shown, the leading edge sweep angle θ lead The value range is
[55] ° 75 ° The relationship between the characteristic area of the wing and the characteristic area is: .
[0015] Preferably, the thickness distribution is designed as an airfoil with equal thickness at the wing root and wingtip, and the wing section thickness must be greater than the minimum thickness required to withstand the aerodynamic load. Less than the maximum thickness limited by the aircraft's shape. Where the chord length is l, and the maximum distance between the wing's far-flight curve and near-flight curve is h, as shown... Figure 6 and 7 As shown, the relative thickness is And the value range is [3%, 5%], with the maximum thickness location x. t At the chord length [30%, 45%], the allowable bending stress of the wing is [σ].
[0016] Preferably, the passivation leading edge cross-sectional shape includes the passivation leading edge cross-sectional profile and the prediction of the compensation effect of the passivation leading edge cross-sectional profile on the lift-to-drag ratio performance.
[0017] Preferably, the initial lift-to-drag ratio can be obtained by CFD numerical simulation, where the lift and drag of the airframe are L0 and D0, respectively. Therefore, the initial lift-to-drag ratio of the airframe excluding the pre-compressed airfoil is... .
[0018] Preferably, the target lift-to-drag ratio The known design specifications to be achieved by the pre-compression angle wing aircraft are given, where the angle of attack α2 is the angle of attack corresponding to the optimal outflow performance.
[0019] Preferably, the passivation of the leading edge cross-sectional profile is achieved through indirect passivation, and the leading edge cross-sectional profile is composed of an arc of radius r and two tangent curves, such as... Figure 8 As shown, the radius r ranges from [5mm, 30mm], and the leading edge profile can be represented by a piecewise function. .
[0020] Preferably, the prediction of the compensation effect of the profile on the lift-to-drag ratio performance is achieved by analyzing the changes in lift-to-drag ratio and drag before and after passivation at different angles of attack using CFD numerical simulation. It can be seen that the larger the passivation radius, the greater the drag and the smaller the lift-to-drag ratio. Moreover, as the angle of attack increases, the influence of the passivation radius on the drag gradually decreases.
[0021] Compared with the prior art, the beneficial effects of the present invention are:
[0022] The novel pre-compressed wing of this hypersonic air-breathing aircraft, which balances internal and external flow performance, analyzes a given airframe using numerical simulation of internal flow performance to obtain the angle of attack and required compression angle under optimal internal flow performance, and designs the aerodynamic shape of the wing while meeting the target lift-to-drag ratio.
[0023] This novel pre-compression wing of the hypersonic air-breathing vehicle balances internal and external flow performance. By installing a pre-compression angle on the wing and designing the wing's aerodynamic shape, the hypersonic air-breathing vehicle achieves a high level of internal and external flow performance within the same flight angle of attack range. Attached Figure Description
[0024] Figure 1 This is a flowchart illustrating the design process of the pre-compressed airfoil of the present invention.
[0025] Figure 2 This is a schematic diagram of the wing pre-compression angle of the present invention;
[0026] Figure 3 This is a schematic diagram illustrating the aerodynamic analysis of the wing of the hypersonic air-breathing aircraft of the present invention;
[0027] Figure 4 This is a diagram showing the internal flow performance of the same type of air intake under different angles of attack according to the present invention;
[0028] Figure 5 This is a configuration diagram of the curves of the wing shape of the present invention;
[0029] Figure 6 This is a configuration diagram of the curves of the wing shape of the present invention;
[0030] Figure 7 This is a schematic diagram of the wing bending stress analysis of the present invention;
[0031] Figure 8 This is a schematic diagram of the profile of the indirect passivation leading edge section of the present invention;
[0032] Figure 9This is a diagram showing the internal and external flow performance of the aircraft wing without pre-compression angle.
[0033] Figure 10 This is a diagram showing the internal and external flow performance characteristics of the aircraft wing when a pre-compression angle is installed. Detailed Implementation
[0034] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. The described novel embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments in this invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.
[0035] This invention relates to a novel pre-compressor wing for a hypersonic air-breathing vehicle that balances internal and external flow performance. The specific implementation steps are as follows:
[0036] Step 1: Wing pre-compression angle θ pre-c Design: Based on the given airframe and target lift-to-drag ratio, determine the optimal external flow performance angle of attack α2 for the aircraft; for Ma 6.0, 27km, 0 °~ 8 ° The internal flow performance of the air intake under flight angle of attack conditions is analyzed, such as... Figure 4 As shown, the moving mean and moving standard deviation are calculated using three data points as a moving interval. The optimal interval for calculating the total pressure recovery coefficient σ is (2... ° 4 ° 5 ° The optimal range for the pressure ratio π at the throat is (4). ° 5 ° 6 ° The optimal range for the intake duct capture flow rate m is (4 ° 5 ° 6 ° Taking the intersection of the three intervals, the range of the optimal internal flow performance angle of attack α1 is [4]. ° 5 ° ]; thus the wing precompression angle .
[0037] Step 2: Wing characteristic area design, such as Figure 3 As shown: The pressure and force conditions of the wing are analyzed. The pressure difference between the windward and leeward sides of the wing is (P wind -P lee The Mach number (Ma) can be calculated from the aircraft's extreme flight conditions: given the incoming flow Ma... start Incoming flow pressure P start Deflection angle α2 (i.e., the optimal angle of attack for outflow performance α2), specific heat ratio γ, the leeward side of the wing is an expansion wave flow region, and the static pressure ratio before and after the wave is... According to the expansion wave relation , The Mach number after the expansion wave deflection can be obtained. Then, based on the isentropic relationship, we obtain... The windward side of the wing is the shock wave flow region, with a static pressure ratio before and after the wave of [missing information]. According to the oblique shock wave relation Solve the equation to obtain the shock angle β, and then obtain... Based on the static pressure ratio before and after the expansion wave / shock wave, the pressure difference on the windward / leeward side of the wing can be calculated. Subsequently, based on the relationship between the lift / drag increment and the initial lift-drag ratio and the target lift-drag ratio, the expression for the wing characteristic area can be obtained as follows: ;
[0038] Step 3, wing shape design, such as Figure 5 As shown: Establish a rectangular coordinate system with point A as the origin. The curve near the fuselage is the intersection curve AB of the wing and fuselage, which can be represented as: The side profile curve AC of the aircraft can be represented as The trailing edge curve BC can be represented as And the leading edge sweep angle θ lead The value range is
[55] ° 75 ° [, to constrain the shape of the side curve of the aircraft body.]
[0039] Step 4: Wing Section Thickness Design: There is a pressure difference (P) between the windward and leeward sides of the wing. wind -P lee If this is not the case, then a bending stress check of the wing is required to determine the minimum thickness t of the wing to withstand the aerodynamic load. min Find the location where the wing experiences the maximum bending moment, that is, where the distance between the wing's far-from-the-fleet curve and its near-from-the-fleet curve is the greatest. Figure 6 As shown, a perpendicular line is drawn from point C to the fuselage, intersecting at point C1. The stress on the wing is then equivalent to that of a cantilever arm for bending stress verification, as follows: Figure 7 As shown, consider a small width ∆b in the neighborhood of point C, with length h for CC1 and height equal to the minimum wing thickness t. min Then the maximum bending moment is Section bending coefficient Therefore, the allowable bending stress condition of the material must be met at point C1. The allowable bending stress of the wing is [σ]. The minimum thickness of the wing is calculated. ;
[0040] Step 5: Design of the leading edge profile of the wing: The leading edge profile consists of an arc of radius r and two tangent curves, such as... Figure 8As shown, the radius r ranges from [5mm, 30mm], and x t The maximum wing thickness is located at the chord length [30%, 45%]. A Cartesian coordinate system is established with the center point of the maximum wing thickness as the origin. The tangent curve portion can be represented using... This indicates that the arc portion with radius r can be represented by... express.
[0041] Preliminary verification of the wing pre-compression angle design concept described in this patent has been conducted:
[0042] With a design condition of Ma 6.0, 27km, and a wing blunt leading edge radius r = 5mm, the following results were obtained:
[0043] 1) Under the given conditions of airframe and target lift-to-drag ratio, optimize the inlet flow performance to achieve the optimal angle of attack within [0°]. ° 8 ° The internal flow performance within the interval [4] shows a trend of first increasing and then decreasing, and in [4] ° 5 ° Within the [7] interval, the internal flow performance is relatively better, but at this time, the external flow performance is at the angle of attack [7]. ° 8 ° Relatively optimal within the range, such as Figure 9 As shown;
[0044] 2) Under the same flight angle of attack range, the pre-compressed wing and its aerodynamic shape satisfy the optimal state of outflow performance. For the wing, install 3... ° Pre-compression angle, at which point the angle of attack [4] ° 5 ° Within the interval, the internal and external flow performance of the aircraft simultaneously reaches a relatively optimal level, such as... Figure 10 As shown, the integrated design of internal and external flow performance enables the aircraft to have excellent flight characteristics.
[0045] During flight, the internal and external flow performance of an aircraft is closely related to its angle of attack. In most cases, the optimal internal and external flow performance of an aircraft cannot be simultaneously satisfied at a certain angle of attack. Therefore, the internal and external flow performance of an aircraft can be adjusted by installing a pre-compression angle on the wing. The compression angle refers to the angle between the chord line of the wing root airfoil and the fuselage axis when the wing is installed on the fuselage. The pre-compression angle wing achieves the purpose of adjusting the aerodynamic performance of the aircraft without affecting the performance of the air intake by changing the angle between the wing and the fuselage, so that the aircraft meets the requirements of optimal internal and external flow performance and has better flight capability.
[0046] It will be apparent to those skilled in the art that the present invention is not limited to the details of the exemplary embodiments described above, and that the invention can be implemented in other specific forms without departing from its spirit or essential characteristics. Therefore, the embodiments should be considered in all respects as exemplary and non-limiting, and the scope of the invention is defined by the appended claims rather than the foregoing description. Thus, all variations falling within the meaning and scope of equivalents of the claims are intended to be included within the present invention. No reference numerals in the claims should be construed as limiting the scope of the claims.
Claims
1. A design method for a novel pre-compressed wing of a hypersonic air-breathing aircraft, characterized in that, The design method includes the following steps: Step 1: Wing pre-compression angle θ pre-c Design: The formula for calculating the wing precompression angle is as follows: Where α1 is the optimal angle of attack for internal flow performance and α2 is the optimal angle of attack for external flow performance; Step 2: Wing characteristic area S character Design: Based on the pressure difference P on the windward / leeward sides of the wing wind -P lee Given the drag D0, lift L0, target lift-to-drag ratio K1, and optimal outflow performance angle of attack α2, the characteristic area S of the wing is calculated. character ; Step 3: Wing Shape Design: The wing shape is determined by the near-body side curve AB, the far-body side curve AC, and the trailing edge curve BC. The near-body side curve AB is the intersection curve of the wing and fuselage, and its shape is consistent with the curve shape at the corresponding position on the fuselage. The far-body side curve AC and the trailing edge curve BC are determined by the wing's characteristic area S. character Calculated; Step 4: Wing Section Design: The wing section design includes wing section thickness design and wing blunt leading edge profile design; the wing section thickness is greater than the minimum thickness t required to withstand the aerodynamic load. min Less than the maximum thickness t limited by the shape of the aircraft max The wing's blunted leading edge profile is indirectly blunted, and the leading edge profile is composed of an arc with radius r and two tangent curves at the top and bottom. In step two, the wing characteristic area S character The calculation formula is: ； Where K1 is the target lift-to-drag ratio, D0 is the drag of the given airframe, L0 is the lift of the given airframe, and P... wind -P lee Given the pressure difference on the windward / leeward sides of the wing, α2 is the optimal angle of attack for outflow performance. In step three, the near-body side curve AB is adopted. Fitting; the AC curve of the far side of the aircraft body is adopted Fitting is performed; the trailing edge curve BC is adopted. Perform fitting; In step four, the minimum thickness Where h is the maximum distance between the wing's far-flight side profile curve and its near-flight side profile curve, and P... wind -P lee The pressure difference between the windward and leeward sides of the wing is given by [σ], where [σ] represents the allowable bending stress of the wing; the maximum thickness is... Where l is the chord length, This refers to the relative thickness.
2. The design method according to claim 1, characterized in that, In step one, the optimal internal flow performance angle of attack α1 is obtained through CFD numerical simulation. Specifically, the optimal internal flow performance angle of attack α1 is determined by the total pressure recovery coefficient σ, the pressure ratio π at the throat, and the intake capture flow rate m; the optimal external flow performance angle of attack α2 is determined by the initial lift-to-drag ratio K0 and the target lift-to-drag ratio K1.
3. The design method according to claim 1, characterized in that, The drag D0 and lift L0 of the given airframe are obtained through CFD numerical simulation; the pressure difference P on the windward / leeward sides of the wing is... wind -P lee It was calculated using the high Mach number compression-expansion aerodynamic principle.
4. The design method according to claim 3, characterized in that, The pressure difference P on the windward / leeward sides of the wing wind -P lee The specific calculation steps are as follows: Given the incoming Mach number Ma start Incoming flow pressure P start The optimal outflow performance angle of attack α2, specific heat ratio γ, the leeward side of the wing is an expansion wave flow region, and the pre-wave and post-wave static pressure ratio is... According to the expansion wave relation , The Mach number after the expansion wave deflection can be obtained. Then, based on the isentropic relationship, we obtain... The windward side of the wing is the shock wave flow region, with a static pressure ratio before and after the wave of [missing information]. According to the oblique shock wave relation Solve the equation to obtain the shock angle β, and then obtain... Based on the static pressure ratio before and after the expansion wave / shock wave, the pressure difference P on the windward / leeward side of the wing can be obtained. wind -P lee .
5. The design method according to claim 1, characterized in that, In step four, the profile of the wing's blunted leading edge section is represented by a piecewise function, specifically as follows: .