Multidisciplinary coupling based integrated propulsion performance conceptual design method
By employing a multidisciplinary coupled flight-propulsion integrated design method, a joint analysis model of aircraft geometry and aerodynamics and a simulation model of aerospace propulsion system are constructed. This solves the problems of aerodynamic characteristic distortion and performance disconnect in traditional design, realizes the refined design of aircraft and propulsion system, and improves design quality and efficiency.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- TAIHANG NATIONAL LABORATORY
- Filing Date
- 2026-02-09
- Publication Date
- 2026-05-15
AI Technical Summary
In traditional integrated aircraft/engine design, the aerodynamic characteristics of the aircraft are severely distorted, the performance design of the aero-engine is disconnected, and the design potential of the aircraft and aero-propulsion system cannot be explored, resulting in a mismatch between the design results and the actual performance.
A multidisciplinary coupling-based integrated flight and propulsion performance conceptual design method is adopted. By constructing a joint analysis model of the aircraft's geometric-aerodynamic lift and drag characteristics and a component-level aerospace propulsion system performance simulation model, a ternary matching relationship is established between the design takeoff total weight, the takeoff thrust of the propulsion system, and the wing planar area. This achieves refined model integration of the aircraft and propulsion system, and the Newton-Raphson iterative algorithm is used to solve the nonlinear equations.
It improves the realism of aircraft aerodynamic performance, accurately analyzes key flight performance parameters and fuel consumption, enhances the quality of the overall performance design scheme of the flight/thrust system, and realizes the multi-disciplinary synergy between aircraft and aero-engines.
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Figure CN121683048B_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of overall design of aircraft propulsion systems, and more specifically to a conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling. Background Technology
[0002] In traditional integrated aircraft / engine design, the conceptual design method for the flight thrust system is based on the matching of total fuel weight and empty weight. The design process can be summarized as follows: First, an initial design takeoff total weight is established and input into the constraint analysis model. The constraint analysis model calculates the design takeoff thrust-to-weight ratio and the design wing loading, which are then input into the mission analysis model. Based on the calculation results of the instantaneous weight ratio for each flight segment in the mission analysis and the formula for calculating the empty weight ratio, a new design takeoff total weight is calculated and compared with the initial design takeoff total weight for error analysis. If the accuracy requirements are met, the iteration termination condition is reached. Otherwise, an initial design takeoff total weight is updated, and the next round of constraint analysis and mission analysis iterations begins. Through the explanation of the above solution process, it can be seen that the traditional method achieves numerical consistency between the design takeoff total weight in the trial calculation stage and the design completion stage. Finally, based on the design takeoff thrust-to-weight ratio and the design wing loading, the wing planar area and the design takeoff thrust are estimated. Therefore, the traditional aircraft conceptual design process can be described as a univariate matching relationship regarding the design takeoff total weight.
[0003] From the calculation process and the final results, the traditional fuel-to-aircraft empty weight matching method used in aircraft conceptual design has the following drawbacks:
[0004] Defect 1: Severe distortion of aircraft aerodynamic characteristics
[0005] Using only general aerodynamic performance charts to interpolate and estimate the lift-drag characteristics of an aircraft often results in a significant deviation from the true aerodynamic characteristics of the aircraft configuration under study. This makes it difficult for aircraft geometry design to accurately reflect the aircraft's aerodynamic characteristics.
[0006] Defect 2: Disconnect in aircraft engine performance design
[0007] Empirical formulas used to estimate the installed thrust coefficient and fuel consumption rate of a propulsion system often lead to a mismatch between the actual performance of the engine under various operating conditions and the performance required for each flight condition in the initial aircraft concept design, after a given engine thermodynamic cycle design.
[0008] Defect 3: Inability to fully exploit the design potential of aircraft and aero engines
[0009] The fact that there is a strong coupling between the aircraft and its aerospace propulsion system is ignored, which makes it impossible to fully analyze the cumulative impact of the synergistic effect between various disciplines on the performance of the flight / propulsion system. Therefore, it is impossible to give full play to the degree of freedom of conceptual design to improve the quality of the flight / propulsion design scheme and achieve the optimal overall performance.
[0010] To address the problem of the disconnect between the overall performance design of aircraft and aero-engines caused by the traditional concept design method of using the fuel-to-aircraft weight matching method for flight-thrust systems, this invention proposes a feasible and efficient new integrated flight / thrust performance concept design method based on multidisciplinary coupling. Through multidisciplinary decoupling processing and collaborative relationship management methods using ternary matching, this invention aims to fully utilize the multidisciplinary synergistic effects in complex systems to solve the global performance design scheme. Summary of the Invention
[0011] In view of this, embodiments of the present invention provide a multidisciplinary coupled flight-propulsion integrated performance conceptual design method. This invention is a multidisciplinary coupled flight-propulsion integrated matching design method based on the ternary matching relationship of aircraft / propulsion system performance, which includes total takeoff weight, takeoff thrust of the propulsion system, and wing planar area. This matching relationship effectively integrates the disciplines of aircraft geometry-aerodynamic joint analysis and design, and propulsion system overall performance design, into the aircraft conceptual design process based on flight profile performance requirements. It overcomes the limitation of the traditional flight-engine integrated design method, which only applies typical aerodynamic characteristics and empirical formulas for propulsion performance to complete the univariate matching of the design takeoff weight between the trial calculation stage and the design completion stage.
[0012] This invention provides the following technical solution: a conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling, comprising:
[0013] S1. Construct a joint analysis model of aircraft geometry-aerodynamic lift-drag characteristics: Use different types of geometric bodies to replace the geometric features of various aircraft components to construct the joint analysis model of aircraft geometry-aerodynamic lift-drag characteristics;
[0014] S2. Construct a component-level aerospace propulsion system performance simulation model: Based on the principles of aerodynamics and thermodynamics, establish performance calculation modules for each component of the engine. By constructing and solving the implicit nonlinear equations characterizing the working relationship of each component, obtain the overall non-installed performance parameters of the engine. Then, combine the drag and flow loss of the aircraft air intake and nozzle to correct the overall non-installed performance parameters of the engine, and calculate the installed thrust and installed fuel consumption rate of the propulsion system.
[0015] S3. Establish and solve the ternary matching relationship for the integrated flight-thrust design: The ternary matching relationship consists of the design takeoff total weight matching equation, the design takeoff thrust matching equation, and the wing planar area matching equation; the design takeoff total weight matching equation matches the initial value of the aircraft total weight with the design takeoff total weight obtained based on mission analysis; the design takeoff thrust matching equation matches the takeoff thrust calculated by the aerospace propulsion system performance simulation model with the takeoff thrust required by the aircraft conceptual design; the wing planar area matching equation matches the wing planar area corresponding to the joint analysis model with the wing planar area required by the aircraft conceptual design.
[0016] The design takeoff total weight matching equation, the design takeoff thrust matching equation, and the wing planar area matching equation are combined to form a nonlinear equation system. Solving this nonlinear equation system yields the matching design parameters of the aircraft and propulsion system that meet the performance requirements of the flight mission.
[0017] According to one embodiment of the present invention, different types of geometric bodies are used to replace the geometric features of various aircraft components to construct the joint analysis model of aircraft geometry-aerodynamic lift and drag characteristics, including:
[0018] Different types of geometric bodies are used to replace the geometric features of various aircraft components. Based on the geometric features of the geometric bodies, the total wetted area of the aircraft and the plane geometric parameters of the wings are calculated. The parasitic drag coefficient, shock wave drag coefficient, induced drag factor and zero lift drag coefficient of the aircraft are calculated in conjunction with these geometric bodies. Then, a polar curve equation describing the relationship between the lift coefficient and the drag coefficient of the aircraft is established.
[0019] According to one embodiment of the present invention, the geometry includes a flat plate, a cylinder, and a cone; the total wetted area of the aircraft is obtained by summing the wetted areas of each geometric component and subtracting the overlapping area between the components; the parasitic drag coefficient is obtained based on the total wetted area of the aircraft and the equivalent friction drag coefficient.
[0020] According to one embodiment of the present invention, the induced drag factor is calculated in relation to the wing aspect ratio and the wing leading edge sweep angle; the shock wave drag coefficient is calculated in relation to the ratio of the maximum cross-sectional area of the fuselage to the total length of the fuselage.
[0021] According to one embodiment of the present invention, the overall performance parameters of the engine not installed are corrected by combining the resistance and flow loss of the aircraft air intake and nozzle, including: correcting the engine inlet conditions and fuel flow rate based on the total pressure loss of the flow inside the aircraft air intake, and deducting the resistance term generated by the total resistance of the aircraft air intake, the internal resistance of the nozzle and the external resistance of the nozzle on the thrust of the propulsion system.
[0022] According to one embodiment of the present invention, constructing a system of nonlinear equations further includes:
[0023] A set of variables, including the parameters of the joint analysis model, the parameters of the aerospace propulsion system performance simulation model, and the aircraft conceptual design parameters, is selected as the matching relation quantity. The Newton-Raphson iterative algorithm is used to solve the nonlinear equation set to obtain the matching design parameters of the aircraft and propulsion system that meet the performance requirements of the flight mission.
[0024] According to one embodiment of the present invention, the parameters constituting the matching relationship quantity include: wingspan, engine design point airflow, and initial design takeoff weight.
[0025] According to one embodiment of the present invention, when solving the nonlinear equation system using the Newton-Raphson iterative algorithm, the matching relation quantity is updated in each iteration according to the following formula:
[0026]
[0027] Where k represents the number of iterations, M J Let be the Jacobian matrix containing information about the first derivative of the residual function F with respect to the matching relation ψ.
[0028] Compared with existing technologies, the beneficial effects achieved by at least one of the above-mentioned technical solutions adopted in the embodiments of this specification include at least the following: Compared with the traditional unidimensional performance matching relationship that only considers the design takeoff weight, the novel flight / thrust ternary performance matching relationship of the embodiments of this invention can effectively integrate the refined models of the aircraft and propulsion system, namely, the aircraft model based on geometric approximation and the component-level aero-engine steady-state performance simulation model, into the aircraft conceptual design process. This realizes the joint analysis and design of aircraft wing geometry / aerodynamic characteristics, aero-engine thermodynamic cycle design, and aircraft conceptual design, thereby making the aerodynamic performance closer to the real situation of the aerodynamic performance of the aircraft configuration under study, and analyzing the key flight performance parameters and fuel consumption of the aircraft over the entire flight range based on more accurate propulsion performance. At the same time, it enables the aircraft conceptual design process to have the ability to improve the quality of the aircraft conceptual design by leveraging the design freedom of the aircraft and aero-engine, and has the potential to obtain the optimal design scheme for the overall performance of the flight / thrust system through the multidisciplinary synergy of the aircraft and aero-engine in the conceptual design process, thus effectively expanding the depth and breadth of integrated flight / thrust performance matching design. Attached Figure Description
[0029] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.
[0030] Figure 1(a) and Figure 1(b) are schematic diagrams of the geometric approximation replacement models of various components of the aircraft in the embodiments of the present invention; wherein, Figure 1(a) is the first diagram and Figure 1(b) is the second diagram;
[0031] Figure 2 This is a diagram of the component-level turbofan engine steady-state performance calculation model according to an embodiment of the present invention;
[0032] Figure 3 This is a flowchart of the design method according to an embodiment of the present invention;
[0033] Figure 4(a) is the actual aerodynamic shape of the F-16C studied in the embodiment of the present invention, and Figure 4(b) is the geometric approximation of the aerodynamic shape of the F-16C studied in the embodiment of the present invention.
[0034] Figure 5(a) is a schematic diagram of the zero-lift drag coefficient characteristics in the aerodynamic performance prediction results of F-16C based on the geometric approximation substitution model in the embodiment of the present invention; Figure 5(b) is a schematic diagram of the induced drag factor in the aerodynamic performance prediction results of F-16C based on the geometric approximation substitution model in the embodiment of the present invention.
[0035] Figure 6(a) is a schematic diagram of aircraft wing geometry design example 1 in an embodiment of the present invention, and Figure 6(b) is a schematic diagram of aircraft wing geometry design example 2 in an embodiment of the present invention.
[0036] Figure 7(a) is the temperature entropy diagram of design 1 of the turbofan engine thermodynamic cycle design example in the embodiment of the present invention, and Figure 7(b) is the temperature entropy diagram of design 2 of the turbofan engine thermodynamic cycle design example in the embodiment of the present invention.
[0037] Among them, 1-wing, 2-leading slat, 3-horizontal tail, 4-vertical tail, 5-ventral fin, 6-fuselage, 7-nozzle, 8-fuselage side, 9-engine nacelle, 10-radar radome, 11-pilot's cockpit. Detailed Implementation
[0038] The embodiments of this application will now be described in detail with reference to the accompanying drawings.
[0039] The following specific examples illustrate the implementation of this application. Those skilled in the art can easily understand other advantages and effects of this application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of them. This application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this application. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.
[0040] like Figure 3 As shown, this embodiment of the invention provides a conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling, including:
[0041] S1. Construct a joint analysis model of aircraft geometry-aerodynamic lift-drag characteristics: Use different types of geometric bodies to replace the geometric features of various aircraft components to construct the joint analysis model of aircraft geometry-aerodynamic lift-drag characteristics;
[0042] S2. Construct a component-level aerospace propulsion system performance simulation model: Based on the principles of aerodynamics and thermodynamics, establish performance calculation modules for each component of the engine. By constructing and solving the implicit nonlinear equations characterizing the working relationship of each component, obtain the overall non-installed performance parameters of the engine. Then, combine the drag and flow loss of the aircraft air intake and nozzle to correct the overall non-installed performance parameters of the engine, and calculate the installed thrust and installed fuel consumption rate of the propulsion system.
[0043] S3. Establish and solve the ternary matching relationship for the integrated flight-thrust design: The ternary matching relationship consists of the design takeoff total weight matching equation, the design takeoff thrust matching equation, and the wing planar area matching equation; the design takeoff total weight matching equation matches the initial value of the aircraft total weight with the design takeoff total weight obtained based on mission analysis; the design takeoff thrust matching equation matches the takeoff thrust calculated by the aviation propulsion system performance simulation model with the takeoff thrust required by the aircraft conceptual design; the wing planar area matching equation matches the wing planar area corresponding to the joint analysis model with the wing planar area required by the aircraft conceptual design.
[0044] The design takeoff total weight matching equation, the design takeoff thrust matching equation, and the wing planar area matching equation are combined to form a nonlinear equation system. Solving this nonlinear equation system yields the matching design parameters of the aircraft and propulsion system that meet the performance requirements of the flight mission.
[0045] Compared to the traditional univariate performance matching relationship that only considers the total takeoff weight, the flight-thrust integrated matching design method proposed in this invention, based on multidisciplinary coupling, can effectively integrate refined flight / thrust models, namely aircraft models based on geometric approximation and component-level aero-engine steady-state performance simulation models, into the aircraft conceptual design process. This enables joint analysis and design of the aircraft's preliminary geometric design and aerodynamic characteristics, as well as the aero-engine's thermodynamic cycle design, with the aircraft conceptual design. As a result, the aerodynamic performance is closer to the actual aerodynamic performance of the aircraft configuration under study, and the key flight performance parameters and fuel consumption over the entire flight can be analyzed based on more accurate propulsion performance, laying the foundation for the global performance design optimization of the flight-thrust system.
[0046] In some embodiments, the flight-thrust integrated performance conceptual design method based on multidisciplinary coupling of the present invention mainly includes the following steps:
[0047] Step 1: Construct a joint analysis model of the aircraft's geometry and aerodynamic lift-drag characteristics.
[0048] It is impractical to establish a precise aerodynamic shape model of an aircraft in the early stages of aircraft conceptual design. This invention proposes to approximate the geometric characteristics of various aircraft components with simple geometric bodies of different types, and puts forward a joint analysis method that links the preliminary geometric design of the aircraft with its lift and drag aerodynamic performance. This method can establish a novel aircraft model with low computational cost and simple modeling, suitable for the aircraft conceptual design stage. This model can reliably represent the mapping relationship between the aircraft's geometric design and its flight lift and drag characteristics, realizing the joint analysis of aerodynamic characteristics and geometric design schemes. As shown in Figures 1(a) and 1(b).
[0049] The geometric features of an aircraft (such as wing dimensions and sweep angle) based on approximate substitutions of various geometric shapes are readily available, and the calculation of the external surface area of each component is very convenient. After summing the areas of all components and subtracting the areas of overlapping parts, the total wetted area S of the aircraft can be obtained. wet Among them, S wet The parasitic drag coefficient C in the analysis of aircraft aerodynamic lift and drag characteristics is... Dp The key is that, based on the above geometric calculation results, C can be... Dp Related to aircraft geometry and obtained through the following calculation method:
[0050] (1)
[0051] In the formula, Characteristic shape factors that approximate the geometry of various aircraft components; This refers to the wetted area of each component of the aircraft. S W,ref It refers to the planar area of the reference wing; C feIt is the equivalent frictional resistance coefficient; i This indicates the serial number of each component of the aircraft.
[0052] In the integrated flight and propulsion concept design, lift-drag characteristic analysis is based on the aircraft polar curve equation. The aircraft polar curve equation can be expressed as:
[0053] (2)
[0054] In the formula, C D C is the drag coefficient; D0 Zero lift-drag coefficient; C L C is the lift coefficient; Dw C is the shock wave drag coefficient; Dp is the parasitic drag coefficient; k1 is the induced drag factor. Here, k1 will no longer be estimated using empirical data, but will be calculated in conjunction with the aircraft wing plane geometry parameters, and can be expressed as:
[0055] (3)
[0056] In the formula, AR is the wing aspect ratio; Λ LE It is the leading edge sweep angle of the wing; Ma is the flight Mach number.
[0057] Zero lift drag coefficient C D0 It is then further decomposed into the shock wave drag coefficient C Dw and parasitic drag coefficient C Dp C Dw This describes the intensity of pressure drag generated when a supersonic airflow, after passing through the bow shock wave at the nose and then accelerating again to supersonic speeds over the curved surfaces of the aircraft fuselage. The numerical calculation method can be correlated with the aircraft's geometric characteristics and is expressed as follows:
[0058] (4)
[0059] In the formula, S W ζ is the wing planar area; E is the ratio of the maximum cross-sectional area of the fuselage to the total length of the aircraft fuselage. WD It is the empirical wave impedance efficiency coefficient, and its value is generally around 2.
[0060] Parasitic drag coefficient C Dp The frictional drag, which characterizes the surface shear force caused by airflow viscosity when airflow passes over the aircraft surface, is defined as follows:
[0061] (5)
[0062] In the formula, S wetThe total wetted area is defined as the sum of the surfaces of all aircraft components through which the incoming airflow passes, and can be calculated; C fe It is the equivalent frictional resistance coefficient, which can be calculated using the Prandtl-Schlichting formula:
[0063] (6)
[0064] In the formula, R eL It is the Reynolds number relating to the characteristic dimensions of an object.
[0065] Step 2: Construct a component-level performance simulation model for aerospace propulsion systems
[0066] This invention conducts modeling research on aero-engine performance simulation models and establishes a high-precision propulsion system installation performance calculation model suitable for airborne installation. For example... Figure 2 As shown. Figure 2 middle, P t The total pressure of the airflow. T t The total temperature of the airflow. P t and T t The subscript numbers represent the numbers of the inlet and outlet sections of each major component.
[0067] Based on the principles of aerodynamics and thermodynamics, a performance calculation module for each major component of the engine was established. Utilizing the common operating conditions of each component, an implicit nonlinear equation system was established, consisting of multiple equilibrium equations characterizing the continuity of flow, power balance, and static pressure balance of the relevant components. Given the control laws and flight operating conditions, the common operating relationships of each component can be solved, allowing for the further calculation of the aerodynamic and thermodynamic parameters of each component, and ultimately obtaining the overall non-installed performance parameters of the engine.
[0068] After being installed on an aircraft, the aero-engine, together with the intake and exhaust systems, forms the propulsion system. Therefore, the drag caused by the viscous, non-ideal flow of air in the intake and exhaust systems must be considered to obtain actual propulsion performance consistent with real-world airborne conditions. Simultaneously, the non-ideal flow of air within the aircraft intake duct causes total pressure loss, which must also be considered. The above impact analysis requires correcting the non-installation performance calculation results of the aero-engine using an installed performance calculation model, considering external intake drag, internal nozzle drag, and external drag, ultimately deriving installed performance calculation results applicable to airborne conditions. By combining the above aircraft intake / exhaust performance loss calculation methods to correct the non-installation performance calculation results, the installed thrust T and installed fuel consumption rate SFC of the aero-engine propulsion system can be calculated as follows:
[0069] (7)
[0070] (8)
[0071] In the formula, C Fg Q is the nozzle thrust coefficient; g,9 V9 is the nozzle exit flow rate; P is the nozzle exhaust velocity; s9 P is the static pressure at the nozzle exit. amb A9 is the atmospheric back pressure at the nozzle exit; Q is the nozzle exit area; a,R V0 is the corrected engine inlet flow rate; V9 is the engine inlet airflow velocity; X is the engine outlet airflow velocity; Inlet X Nozzle These are the total resistance of the air intake and the external flow resistance of the nozzle, respectively. Q f,A The engine fuel flow rate is adjusted to take into account the internal flow characteristics of the intake manifold.
[0072] Step 3: Triad Matching Relationship in Multidisciplinary Coupled Flight-Thrust Integrated Design
[0073] Based on the univariate matching relationship for takeoff weight design, and using a joint analysis model of preliminary aircraft geometry and aerodynamic performance, as well as a high-precision simulation model of aerospace propulsion system component-level performance, this invention proposes a novel ternary matching relationship involving multiple disciplines. The specific execution strategy for realizing this novel matching relationship can be summarized as follows: within a reasonable design range, given freely given geometric design parameters for some parts of the aircraft (e.g., wing root length, wingtip length, 0.25 times chord sweep angle) and thermodynamic cycle parameters for some parts of the engine (compression component boost ratio, combustion chamber outlet temperature, bypass ratio, etc.), the unique feasible matching design that satisfies the flight mission profile performance requirements is obtained by solving the novel matching relationship described by the ternary relationship.
[0074] The novel ternary matching relationship can be achieved by including flight / push matching relationship variables. X M The three implicit nonlinear equations are represented as follows:
[0075] (1) Design takeoff weight matching
[0076] When achieving integrated performance matching of flight and thrust in aircraft conceptual design, the initial value of the aircraft's total weight W is used in constraint analysis and mission analysis calculations. Tko,ini After the mission analysis is completed, the designed takeoff gross weight W is calculated based on the empty weight ratio and instantaneous mass ratio. Tko Matching:
[0077] (9)
[0078] (2) Design takeoff thrust matching
[0079] When achieving integrated performance matching between flight and thrust in aircraft conceptual design, the takeoff thrust calculated by the propulsion system model based on the selected thermodynamic cycle design scheme under given takeoff conditions and control laws. T Tko Thrust requirements for takeoff as determined by the aircraft concept design T Tko,d Matching:
[0080] (10)
[0081] (3) Wing planar area matching
[0082] When achieving integrated performance matching of flight and thrust in aircraft conceptual design, the wing planar area is approximated by geometrical substitution of the aircraft model. S W wing planar area required in the conceptual design results S W,d Matching:
[0083] (11)
[0084] The three implicit nonlinear equations above can form a system of nonlinear equations:
[0085] (12)
[0086] In the formula, X M This represents the quantity of the fly / push matching relationship.
[0087] This set of equations characterizes the ternary matching relationship between the design takeoff gross weight, takeoff thrust requirement, and wing planar area in the integrated flight / thrust performance matching. It can be seen that solving the ternary matching relationship between the takeoff gross weight, takeoff thrust, and wing planar area is essentially solving the nonlinear equation set described by equation (12). To ensure that the necessary condition for the equation set to have one and only one solution under a given design condition is that the equation set is full rank, this invention selects three parameters from the aircraft model, propulsion system model, and aircraft conceptual design model to determine the quantities constituting the flight / thrust matching relationship. X M , can be represented as:
[0088] (13)
[0089] In the formula, Lb W Q is the wingspan of the wing; in W is the engine design point intake airflow. Tko,ini This is the initial value for the total takeoff weight.
[0090] Solving the common operating conditions of an aero-engine is essentially solving an implicit system of nonlinear equations:
[0091] (14)
[0092] In the formula, ε represents the iteration termination precision, which is usually set to 10. -5 .
[0093] Solve using the Newton-Raphson algorithm ψ Update in each iteration ψ The algorithm is as follows:
[0094] (15)
[0095] In the formula, M J The Jacobian matrix contains information about the first derivative of the residual pairs with respect to the matching variables; ψ The variable matrix is matched for the joint operation of the aero-engine; k represents the number of the current iteration.
[0096] This invention uses numerical examples to demonstrate the rationality and effectiveness of obtaining the flight / thrust conceptual design results based on the novel integrated flight / thrust performance matching relationship.
[0097] To verify the accuracy of the joint analysis model of aircraft geometry design and aerodynamic characteristics based on simple geometry in predicting the aerodynamic characteristics of real aircraft, this study uses the F-16C as the research object. The approximate replacement model of the F-16C geometry constructed in this invention is shown in Figures 4(a) and 4(b). In Figure 4(a), 370-GALLON FUEL TANK represents a 370-gallon external fuel tank; OUTBOARD STORES LOCATIONS represents the location of the outer hardpoints; SPEEDBRAKEHINGE represents the speed brake hinge; AIM-9 MISSILE represents the AIM-9 missile; MISSILE LAUNCHER represents the missile launcher; and FLAPERON represents the flaperon.
[0098] The characteristic geometric parameters of the F-16 were input into the aircraft's geometric approximation substitution model to predict its aerodynamic characteristics. The obtained zero-lift drag coefficient characteristics and induced drag factor as a function of flight Mach number were compared with the results in the literature, as shown in Figures 5(a) and 5(b). The literature in Figures 5(a) and 5(b) refers to: "Introduction to aeronautics: a design perspective (3rd ed)," Steven A. Brandt, 2004.
[0099] Referring to the performance requirements of typical fighter jet flight mission profiles, and based on a geometric approximation of the flight model, multiple feasible joint design results of aircraft geometric parameters and propulsion system thermodynamic cycle parameters were obtained through the design of novel matching relationships. Except for the wing planar design parameters, all other geometric design parameters of the aircraft geometric approximation model are referenced from the F-16C. The three feasible matching design schemes obtained are listed in Table 1, with one set of results serving as a reference design for comparison with the other two sets of designs (Design 1 and Design 2).
[0100] Table 1. Performance Matching Design Results of Flight / Thrust Integration
[0101]
[0102] The results show that the wing planar area calculated using the aircraft geometric approximation model has a numerical error of no more than 0.01% compared to the required wing planar area in the conceptual design. Regarding the thrust required for takeoff, the thrust calculation results from the propulsion system performance simulation model under given takeoff conditions and control laws also match the takeoff thrust required in the conceptual design, with an error of no more than 0.01%. The conceptual design results demonstrate that the required takeoff thrust and the required total wing area are positively correlated with the designed takeoff gross weight; that is, a larger takeoff gross weight requires a larger takeoff thrust and wing planar area to provide lift, which conforms to objective matching laws.
[0103] The solved aircraft wing geometry design calculations are shown in Figures 6(a) and 6(b). The temperature-entropy diagrams characterizing the thermodynamic cycle design of the turbofan engine at the design point are shown in Figures 7(a) and 7(b). Figures 6(a) and 7(a) represent the engine design results for Design 1 in Table 1, while Figures 6(b) and 7(b) represent the engine design results for Design 2 in Table 1. The reference design refers to the engine design results of the "Reference Design" in Table 1. This completes an implementation case of a multidisciplinary coupled flight-propulsion integrated matching design method.
[0104] The multidisciplinary coupled flight-thrust integrated design ternary matching relationship proposed in this invention can effectively integrate the geometrically approximated aircraft model and the component-level aero-engine steady-state performance simulation model into the aircraft conceptual design process. In the final conceptual design results, the error between the designed wing area and the wing area of the aircraft model, and between the designed takeoff thrust and the thrust of the turbofan engine model under a given takeoff condition, is within 0.01%.
[0105] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.
Claims
1. A conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling, characterized in that, include: S1. Construct a joint analysis model of aircraft geometry-aerodynamic lift-drag characteristics: Use different types of geometric bodies to replace the geometric features of various aircraft components to construct the joint analysis model of aircraft geometry-aerodynamic lift-drag characteristics; S2. Construct a component-level aerospace propulsion system performance simulation model: Based on the principles of aerodynamics and thermodynamics, establish performance calculation modules for each component of the engine. By constructing and solving the implicit nonlinear equations characterizing the working relationship of each component, obtain the overall non-installed performance parameters of the engine. Then, combine the drag and flow loss of the aircraft air intake and nozzle to correct the overall non-installed performance parameters of the engine, and calculate the installed thrust and installed fuel consumption rate of the propulsion system. S3. Establish and solve the ternary matching relationship for the integrated flight-thrust design: The ternary matching relationship consists of the design takeoff total weight matching equation, the design takeoff thrust matching equation, and the wing planar area matching equation; the design takeoff total weight matching equation matches the initial value of the aircraft total weight with the design takeoff total weight obtained based on mission analysis; the design takeoff thrust matching equation matches the takeoff thrust calculated by the aviation propulsion system performance simulation model with the takeoff thrust required by the aircraft conceptual design. The wing planar area matching equation matches the wing planar area in the joint analysis model with the wing planar area required for the aircraft conceptual design. The design takeoff total weight matching equation, the design takeoff thrust matching equation, and the wing planar area matching equation are combined to form a nonlinear equation system. Solving this nonlinear equation system yields the matching design parameters of the aircraft and propulsion system that meet the performance requirements of the flight mission.
2. The conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling as described in claim 1, characterized in that, By using different types of geometric shapes to replace the geometric features of various aircraft components, a joint analysis model of the aircraft's geometry-aerodynamic lift and drag characteristics is constructed, including: Different types of geometric bodies are used to replace the geometric features of various aircraft components. Based on the geometric features of the geometric bodies, the total wetted area of the aircraft and the plane geometric parameters of the wings are calculated. The parasitic drag coefficient, shock wave drag coefficient, induced drag factor and zero lift drag coefficient of the aircraft are calculated in conjunction with these geometric bodies. Then, a polar curve equation describing the relationship between the lift coefficient and the drag coefficient of the aircraft is established.
3. The integrated flight and propulsion performance conceptual design method based on multidisciplinary coupling according to claim 2, characterized in that, The geometry includes flat plates, cylinders, and cones; the total wetted area of the aircraft is obtained by summing the wetted areas of each geometric component and subtracting the overlapping area between components; the parasitic drag coefficient is calculated based on the total wetted area of the aircraft and the equivalent friction drag coefficient.
4. The conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling as described in claim 2, characterized in that, The induced drag factor is calculated in relation to the wing aspect ratio and the wing leading edge sweep angle; the shock wave drag coefficient is calculated in relation to the ratio of the maximum cross-sectional area of the fuselage to the total length of the fuselage.
5. The conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling as described in claim 1, characterized in that, The engine's non-installed overall performance parameters are corrected by combining the drag and flow loss of the aircraft's air intake and nozzle. This includes correcting the engine's inlet conditions and fuel flow rate based on the total pressure loss of the flow inside the aircraft's air intake, and deducting the drag term generated by the total drag of the aircraft's air intake, the internal drag of the nozzle, and the external drag of the nozzle on the thrust of the propulsion system.
6. The conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling as described in claim 1, characterized in that, The system of nonlinear equations also includes: A set of variables, including the parameters of the joint analysis model, the parameters of the aerospace propulsion system performance simulation model, and the aircraft conceptual design parameters, is selected as the matching relation quantity. The Newton-Raphson iterative algorithm is used to solve the nonlinear equation set to obtain the matching design parameters of the aircraft and propulsion system that meet the performance requirements of the flight mission.
7. The conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling as described in claim 6, characterized in that, The parameters constituting the matching relationship include: wingspan, engine design point airflow, and initial design takeoff weight.
8. The conceptual design method for integrated flight and propulsion performance based on multidisciplinary coupling as described in claim 6, characterized in that, When solving the nonlinear equations using the Newton-Raphson iterative algorithm, the matching relation quantity is updated in each iteration according to the following formula: Where k represents the number of iterations, M J Let be the Jacobian matrix containing information about the first derivative of the residual function F with respect to the matching relation ψ.