Two-stage separation point optimization method for energy-climbing combined propulsion launch vehicle

Abstract

The application provides a two-stage orbit entry separation point optimization method based on energy climbing of a combined power carrier, which comprises the following steps: step 1, determining the relationship between the selection of the separation point and the total take-off mass W of the combined power carrier TO ; step 2, establishing the relationship between the specific impulse I SP of the carrier stage vehicle and the fuel consumption rate / amount; step 3, according to the energy height, calculating the relationship between the real-time changing fuel mass of the carrier stage vehicle and the speed v and height h of the carrier stage vehicle flight; step 4, according to the relationship between the fuel mass and the speed v and height h, discretizing the speed and height of the climbing trajectory to obtain a plurality of sets of corresponding relationships between the speed v, height h and fuel mass W F ; and calculating the total take-off mass W TO of each set of relationships, and selecting the speed v and height h corresponding to the minimum value of W TO as the optimal separation point. The application can realize reliable separation of the combined power carrier in the inter-stage separation task.

Description

Technical Field

[0001] This invention belongs to the technical field of two-stage-to-orbit combined propulsion vehicles, and relates to an optimization method for the two-stage-to-orbit separation point of a combined propulsion vehicle based on energy climbing. Background Technology

[0002] With the increasing frequency of international space activities, the demand for space launch vehicles has risen sharply. Against this backdrop, reusable space transportation systems have become a focus of space technology research. Their development is crucial for reducing the cost of space exploration. Their main propulsion methods are rocket propulsion and combined propulsion. Rocket propulsion technology focuses on traditional propulsion methods; combined propulsion technology explores new technological routes that combine rocket propulsion with other propulsion methods to improve the efficiency of reusable space transportation systems.

[0003] The importance of combined-propellant launch vehicles (CPVs) in the field of space transportation cannot be underestimated. They not only provide humanity with a rapid means of access to space but also serve as a fundamental platform for large-scale space development. CPVs have brought technological advancements to the space transportation sector by offering a safe and efficient method of space access. With the increasing international interest in space exploration, the development of CPVs is receiving increasing attention.

[0004] Combined-engine launch vehicles (GEVs) have significant applications in both military and civilian fields. They not only provide rapid global transportation capabilities but also enable diverse space missions, leading many countries to prioritize their development. In the research of two-stage-to-orbit (TTO) GEVs, safely completing the interstage separation process is a critical technical issue. The selection of the separation point not only affects the safety of the separation process but also directly impacts mission planning and execution, ultimately influencing the overall design of the TTO GEV.

[0005] In summary, for combined-powered vehicles, how to select the separation point has become an urgent technical problem to be solved. Summary of the Invention

[0006] The purpose of this invention is to at least solve one of the problems existing in the prior art.

[0007] Therefore, this invention provides an optimization method for the two-stage orbital separation point of a combined propulsion vehicle based on energy climbing, which can achieve reliable separation of the combined propulsion vehicle in interstage separation missions.

[0008] The technical solution of the present invention is as follows:

[0009] An optimization method for the two-stage-to-orbit separation point of a combined-propulsion launch vehicle based on energy climb, wherein the combined-propulsion launch vehicle includes a launch-stage vehicle and a mission-stage vehicle; the separation point includes separation velocity and separation altitude.

[0010] The specific steps of this separation point optimization method are as follows:

[0011] Step 1: Perform mass modeling on the combined-powered launch vehicle to determine the selection of the separation point and the total takeoff mass W of the combined-powered launch vehicle. TO Relationship: The selected separation point ensures that the total takeoff mass W TO Minimum;

[0012] Step 2: Model the propulsion system of the launch vehicle and establish its specific impulse I. SP Fuel consumption rate / fuel consumption Relationship;

[0013] Step 3: Before the mission-class aircraft separates from the launch vehicle, calculate the relationship between the real-time changing fuel mass of the launch vehicle and the speed v and altitude h of the launch vehicle based on the energy altitude.

[0014] Step 4: Based on the real-time changing fuel mass of the launch vehicle and its flight speed v and altitude h, discretize the climb trajectory for speed and altitude to obtain several sets of one-to-one corresponding data on the launch vehicle's flight speed v, altitude h, and fuel mass W. F The relationship between the two sets of relationships was calculated, and the total takeoff mass W for each set of relationships was obtained. TO In all the total takeoff mass W obtained TO In the middle, select W TO The minimum velocity v and height h are taken as the optimal separation point.

[0015] Furthermore, in step 1, the total takeoff mass W of the combined-powered launch vehicle TO Including the structural mass W of the launch vehicle E The fuel mass W carried by the launch vehicle F Maintenance weight W of launch vehicle C The payload mass W of the launch vehicle P :

[0016] W TO =W E +W F +W P +W C (1)

[0017] Among them, the payload mass W of the launch vehicle PThis refers to the total weight of the mission-class aircraft; the payload mass W of the launch vehicle is... P Maintenance weight W of launch vehicle C For the given design requirements, under the given W P and W C Under the condition of; adopting structural mass ratio ratio of fuel mass Total takeoff mass W TO Represented as:

[0018]

[0019] Therefore, the total takeoff mass W of the combined-powered launch vehicle TO The estimation needs to be combined with the structural mass ratio fuel mass ratio Combined-fuel launch vehicles use the separation point as the mission interface; differences in the separation point affect the fuel mass (W) of the launch vehicle. F Therefore, the principle for selecting the separation point is: the selected separation point should ensure that the total takeoff mass W... TO Minimum.

[0020] Furthermore, the value range of the structural mass ratio is 0.2 to 0.75.

[0021] Furthermore, in step 2, the specific impulse I of the launch vehicle... SP Fuel consumption rate / fuel consumption The relationship is:

[0022]

[0023] Where F represents the thrust generated by the propulsion system of the launch vehicle; g represents the gravitational acceleration at sea level.

[0024] Furthermore, the specific impulse I of the launch vehicle SP Represented as:

[0025]

[0026] in, These are five different coefficients: M is the flight Mach number of the launch vehicle, and ∝ is the angle of attack of the launch vehicle.

[0027] Furthermore, the specific steps of step 3 are as follows:

[0028] Step 3-1: Perform a force analysis on the combined-powered launch vehicle. Let the angle of attack α be the angle between the incoming flow velocity and the chord line of the launch vehicle, the climb angle be θ, the thrust F be forward along the chord line, the lift L be perpendicular to the incoming flow direction, the drag D be parallel to the incoming flow direction, the total weight W of the combined-powered launch vehicle be vertically downward, the acceleration of the combined-powered launch vehicle be a, the velocity be V, and the altitude be h. This yields the thrust-to-weight ratio parameter.

[0029]

[0030] Step 3-2, calculate the acceleration 'a' of the launch vehicle during the climb phase:

[0031]

[0032] Where V1 is the velocity at the intersection of the previous level isoenergy altitude line and the climbing curve, and V2 is the velocity at the intersection of the next level isoenergy altitude line and the climbing curve; h1 is the altitude at the intersection of the previous level isoenergy altitude line and the climbing curve, and h2 is the altitude at the intersection of the next level isoenergy altitude line and the climbing curve.

[0033] Substituting formula (23) into formula (18), the thrust-to-weight ratio parameter can be calculated.

[0034] Step 3-3: Calculate the fuel consumption mass Δm of the launch vehicle during the climb phase. f :

[0035]

[0036] Wherein, energy conversion efficiency η; He is the energy level. He i For the i-th energy level, He i+1 The energy height of level i+1; Fuel consumption rate / fuel consumption; P S Residual power per unit mass; I SP Specific impulse for launch vehicles;

[0037] The thrust-to-weight ratio parameter is obtained from step 3-2. Substituting into formula (34), the energy height can be calculated from He. i Change to He i+1 When the mass change is equal to the mass of fuel consumed during that stage, Δm f ;

[0038] The fuel consumption mass between each two adjacent isoenergy altitude lines can be iteratively calculated. Summing up the fuel consumption mass for each segment yields the total fuel consumption for hypersonic climb, i.e., SDm. f (i)=WF This allows us to obtain the relationship between the real-time changing fuel mass of the launch vehicle and its flight speed v and altitude h.

[0039] Furthermore, the specific process of step 3-1 is as follows:

[0040] According to Newton's second law, force analysis of the combined-powered vehicle along the body direction yields the following results:

[0041]

[0042] Force analysis of the combined propulsion vehicle in the direction perpendicular to the body:

[0043] L=Fsinq+Wcosq=Fsina+Wcosq (14)

[0044] Combining equations (13) and (14), we can obtain:

[0045]

[0046] Since α≈θ, formula (15) can be transformed into:

[0047]

[0048] cosa≈1, sinα≈0, formula (16) can be transformed into:

[0049]

[0050] Further simplification yields:

[0051]

[0052] Furthermore, the specific process of step 3-2 is as follows:

[0053] The displacement component of the climb segment along the body direction can be expressed as:

[0054]

[0055] By deriving formula (19), we can obtain:

[0056]

[0057] Further simplification yields:

[0058]

[0059] Since the displacement component of the climb segment parallel to the body direction can also be expressed as:

[0060]

[0061] Therefore, the acceleration 'a' during the ascent phase can be expressed as:

[0062]

[0063] Furthermore, the specific process of step 3-3 is as follows:

[0064] Multiplying both sides of equation (13) by the velocity V, we get:

[0065]

[0066] Wherein, the vertical component of velocity V is the climb velocity, which can be expressed as:

[0067]

[0068] Combining equations (24) and (25), we can obtain:

[0069]

[0070] In equation (26), the right side of the equation represents the residual power per unit mass, P. S , can be represented as follows:

[0071]

[0072] Combining equations (26) and (27), we can obtain:

[0073]

[0074] Due to the climb angle q and the unit residual power P S The coupling relationship can be expressed as:

[0075]

[0076] According to equations (28) and (29), we can obtain:

[0077]

[0078] The energy gain per unit mass of fuel consumed during the climb of a combined-power vehicle is defined as the energy conversion efficiency η, which can be expressed by the formula:

[0079]

[0080] Energy level is determined by He i Change to He i+1 When the mass change is equal to the mass of fuel consumed during that stage, Δm f It can be obtained by integrating the following formula:

[0081]

[0082] The relationship between aircraft thrust F and specific impulse can be expressed as:

[0083]

[0084] Substituting equations (31), (33), and (29) into equation (32) yields equation (34).

[0085] Furthermore, the fuel consumed by the combined-powered vehicle from horizontal takeoff to the start of hypersonic climb accounts for 5% of the total takeoff mass of the combined-powered vehicle, therefore W in equation (34) i The initial value W0 = 0.95 * W TO .

[0086] By applying the above technical solution, the present invention has the following beneficial effects:

[0087] (1) This invention takes a two-stage-to-orbit combined propulsion vehicle as the research object. Starting from the design scheme of traditional aircraft, after integrating the mission indicators of the two-stage-to-orbit combined propulsion vehicle, the interstage separation scheme is designed with the minimum total takeoff mass as the performance index. That is, the selection of the optimal separation point is based on the fuel consumption mass during the climb phase as the starting point for calculating the total takeoff mass. The total takeoff mass is calculated, and the total takeoff mass calculation results at a specific separation altitude and Mach number can be obtained. The reliable selection of the separation point is realized in the conceptual design stage, which provides theoretical support for the selection of separation strategy of two-stage-to-orbit combined propulsion vehicle.

[0088] (2) By introducing the concept of energy altitude, this invention realizes detailed calculation of fuel mass, provides selection method and calculation method of interstage separation point, forms a complete interstage separation strategy, finds the relationship between separation point and total takeoff mass, and can realize rapid calculation of total takeoff mass and separation point in the preliminary design stage of aircraft. Attached Figure Description

[0089] The accompanying drawings, which form part of this specification, are provided to further illustrate embodiments of the invention and, together with the textual description, explain the principles of the invention. It is obvious that the drawings described below are merely some embodiments of the invention, and those skilled in the art can obtain other drawings based on these drawings without any creative effort.

[0090] Figure 1 This is a schematic diagram of a combined propulsion vehicle.

[0091] Figure 2 This is a flowchart of the present invention;

[0092] Figure 3 Force analysis diagram of combined propulsion vehicle;

[0093] Figure 4 Design structure matrix diagram for total takeoff mass (0:1:2 in the diagram: design sequence);

[0094] Figure 5 This is a schematic diagram of isoenergy height lines;

[0095] Figure 6 This is a diagram showing the intersection of isoenergy elevation lines and ascent curves;

[0096] Figure 7 A graph showing the relationship between the residual mass ratio and total mass versus velocity of a combined-powered vehicle.

[0097] Figure 8 A schematic diagram of the window for inter-level separation optimization;

[0098] Figure 9 A graph showing the influence of structural mass ratio on separation conditions and takeoff mass;

[0099] Figure 10 A diagram illustrating the impact of specific impulse on separation conditions and takeoff mass of a mission-class aircraft.

[0100] Figure 11 This is a diagram illustrating the impact of the specific impulse of a launch vehicle on separation conditions and takeoff mass. Detailed Implementation

[0101] It should be noted that, unless otherwise specified, the embodiments and features described in this application can be combined with each other. The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of them. The following description of at least one exemplary embodiment is merely illustrative and is in no way intended to limit the present invention or its application or use. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0102] It should be noted that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to limit the exemplary embodiments according to this application. As used herein, the singular form is intended to include the plural form as well, unless the context clearly indicates otherwise. Furthermore, it should be understood that when the terms "comprising" and / or "including" are used in this specification, they indicate the presence of features, steps, operations, devices, components, and / or combinations thereof.

[0103] Unless otherwise specifically stated, the relative arrangement, numerical expressions, and values ​​of the components and steps set forth in these embodiments do not limit the scope of the invention. It should also be understood that, for ease of description, the dimensions of the various parts shown in the drawings are not drawn to actual scale. Techniques, methods, and devices known to those skilled in the art may not be discussed in detail, but where appropriate, such techniques, methods, and devices should be considered part of the specification. In all examples shown and discussed herein, any specific values ​​should be interpreted as merely exemplary and not as limitations. Therefore, other examples of exemplary embodiments may have different values. It should be noted that similar reference numerals and letters in the following figures denote similar items; therefore, once an item is defined in one figure, it need not be further discussed in subsequent figures.

[0104] Example 1:

[0105] This embodiment provides an optimization method for the two-stage-to-orbit separation point of a combined-propulsion launch vehicle based on energy climb. The combined-propulsion launch vehicle includes a first-stage (launch stage) vehicle and a second-stage (mission stage) vehicle; see appendix. Figure 1 The launch vehicle uses a combined ramjet propulsion system, while the mission vehicle uses rocket propulsion. Before separation, the mission vehicle serves as the payload of the launch vehicle. Based on the general flight mission process, the launch vehicle can be divided into five basic phases: takeoff, climb, cruise (the main phase of the mission), descent, and landing. The launch vehicle carries the mission vehicle during takeoff and climb until it reaches the vicinity of the separation point. During cruise, it searches for the optimal separation point, at which point the mission vehicle separates from the launch vehicle. The mission vehicle continues its orbital insertion mission, while the launch vehicle descends and lands. Before separation, only the launch vehicle consumes fuel; therefore, the fuel mass of the launch vehicle gradually decreases. The separation point includes both separation velocity and separation altitude.

[0106] See appendix Figure 2 The specific steps of this separation point optimization method are as follows:

[0107] Step 1: Perform mass modeling on the combined-powered launch vehicle to determine the optimal selection of the separation point and the total takeoff mass W of the combined-powered launch vehicle. TO Relationship:

[0108] Let the total weight of the combined-powered vehicle be W. Due to fuel consumption during flight, the total weight W gradually decreases. Therefore, the total weight W at takeoff (i.e., the takeoff weight) is equal to the total takeoff mass W of the combined-powered vehicle. TO Total takeoff mass W TO Including the structural mass W of the launch vehicle E The fuel mass W carried by the launch vehicle F Maintenance weight W of launch vehicle C The payload mass W of the launch vehicle P :

[0109] W TO =W E +W F +W P +W C (1)

[0110] Among them, the payload mass W of the launch vehicle P This refers to the total weight of the mission-class aircraft; the payload mass W of the launch vehicle is... P Maintenance weight W of launch vehicle C Given the design requirements, and the total takeoff mass W of the combined-powered launch vehicle... TO Irrelevant;

[0111] Therefore, given W P and W C Under the condition of, structural mass W E Fuel quality W F The total takeoff mass W of the combined-powered launch vehicle TO Design parameters; structural mass ratio ratio of fuel mass The total takeoff mass W can be TO Represented as:

[0112]

[0113] According to the above formula, the structural mass ratio ratio of fuel mass The sum must satisfy the constraint that it is less than 1; the total takeoff mass W obtained when the sum of the two is greater than 1 is... TO A negative value indicates an excessively high structural mass ratio or fuel mass ratio. In this case, it is necessary to weigh the structural design weights related to the structural mass ratio. The structural mass ratio generally ranges between 0.2 and 0.75, but the range varies for different types of launch vehicles. Based on practical experience, the structural mass ratio of a typical launch vehicle changes linearly.

[0114] Fuel-to-mass ratio refers to the mass of fuel (W) that a launch vehicle needs to carry to complete its mission. F W% of total takeoff mass TO The ratio of W to W; therefore, fuel mass W F The design focuses on the fuel consumption of the launch vehicle during the mission phase; the fuel consumption during the climb phase ensures that the launch vehicle can reach the minimum total takeoff mass W at the separation point. TO ;

[0115] In summary, the total takeoff mass W of the combined-powered launch vehicle TO The estimation needs to be combined with the structural mass ratio fuel mass ratio Unlike traditional single-stage-to-orbit vehicles, combined-fuel launch vehicles use the separation point as the mission interface; differences in the separation point affect the fuel mass (W) of the launch vehicle. F Therefore, it is necessary to comprehensively consider the effect of the separation point on the total takeoff mass W of the combined-powered launch vehicle. TO The impact of this. Total takeoff mass W TO Design structure matrix such as Figure 4 As shown; Figure 4 In the middle, based on separation conditions and total takeoff mass W TO As a design parameter.

[0116] Therefore, the principle for selecting the separation point is: the selected separation point should ensure that the total takeoff mass W TO At the minimum, for example, if the separation altitude of the separation point is chosen too high, it will result in more fuel being needed to support the launch vehicle to fly to the separation point. In this case, the total takeoff mass W TO It will be relatively high.

[0117] Step 2: Model the propulsion system of the launch vehicle, that is, establish the relationship between the specific impulse of the launch vehicle and the fuel consumption rate / fuel consumption:

[0118] For the propulsion system of a launch vehicle, specific impulse I is often defined. SP As a performance indicator of propulsion system efficiency, a higher specific impulse indicates a higher efficiency in converting thermal energy into mechanical energy, and thus better propulsion system performance. Specific impulse I SP This represents the thrust generated per unit mass of fuel consumed, measured in seconds (s), and has the dimension of time. The expression is as follows:

[0119]

[0120] Where F represents the thrust generated by the propulsion system of the launch vehicle; Fuel consumption rate per second / fuel consumption of the launch vehicle, where g represents the gravitational acceleration at sea level;

[0121] Because the launch vehicle's propulsion system uses a ramjet propulsion system with liquid hydrogen as fuel, the specific impulse I of the launch vehicle is... SP It can be represented as:

[0122]

[0123] in, These are five different coefficients, where M is the flight Mach number of the launch vehicle, and ∝ is the angle of attack of the launch vehicle.

[0124] Step 3: Before the mission-class aircraft separates from the launch vehicle, calculate the relationship between the real-time changing fuel mass of the launch vehicle and the speed v and altitude h of the launch vehicle based on the energy altitude.

[0125] Step 3-1: Perform a force analysis on the combined propulsion vehicle to obtain the thrust-to-weight ratio parameter.

[0126] See appendix Figure 3 In the force analysis of the combined-powered launch vehicle, it is assumed that the total mass is concentrated at a point mass; the angle of attack α is the angle between the incoming flow velocity and the chord line of the launch vehicle, the climb angle is θ, the thrust F is forward along the chord line, the lift L is perpendicular to the incoming flow direction, the drag D is parallel to the incoming flow direction, the total weight W of the combined-powered launch vehicle is vertically downward, the acceleration of the combined-powered launch vehicle is a, the velocity is V, and the altitude is h.

[0127] According to Newton's second law, force analysis of the combined-powered vehicle along the body direction yields the following results:

[0128]

[0129] Force analysis of the combined propulsion vehicle in the direction perpendicular to the body:

[0130] L=Fsinq+Wcosq=Fsina+Wcosq (14)

[0131] Combining equations (13) and (14), we can obtain:

[0132]

[0133] Since α≈θ, formula (15) can be transformed into:

[0134]

[0135] Since the angle of attack α is relatively small during the climb of the combined propulsion vehicle, it is assumed that cosa≈1 and sinα≈0, and formula (16) is transformed into:

[0136]

[0137] Further simplification yields:

[0138]

[0139] Step 3-2, calculate the acceleration 'a' of the launch vehicle during the climb phase:

[0140] According to the literature "Aerodynamics for engineers", the displacement component of the ascent segment along the body direction can be expressed as:

[0141]

[0142] By deriving formula (19), we can obtain:

[0143]

[0144] Further simplification yields:

[0145]

[0146] Where V1 is the velocity at the intersection of the previous level's isoenergy altitude line and the climb curve, and V2 is the velocity at the intersection of the next level's isoenergy altitude line and the climb curve; as shown in the attached figure. Figure 5 and Figure 6 As shown;

[0147] Since the displacement component of the climb segment parallel to the body direction can also be expressed as:

[0148]

[0149] Where h1 is the altitude at the intersection of the previous level isoenergy elevation line and the ascent curve, and h2 is the altitude at the intersection of the next level isoenergy elevation line and the ascent curve; as shown in the attached figure. Figure 5 and Figure 6 As shown;

[0150] Therefore, the acceleration 'a' during the ascent phase can be expressed as:

[0151]

[0152] Substituting formula (23) into formula (18), the thrust-to-weight ratio parameter can be calculated.

[0153] Step 3-3: Calculate the fuel consumption mass Δm of the launch vehicle during the climb phase. f :

[0154] Multiplying both sides of equation (13) by the velocity V, we get:

[0155]

[0156] Wherein, the vertical component of velocity V is the climb velocity, which can be expressed as:

[0157]

[0158] Combining equations (24) and (25), we can obtain:

[0159]

[0160] Where He represents the energy level, as shown in the attached diagram. Figure 5 and Figure 6 As shown;

[0161] In equation (26), the right side of the equation represents the residual power per unit mass, P. S , can be represented as follows:

[0162]

[0163] Combining equations (26) and (27), we can obtain:

[0164]

[0165] According to the literature "Research on Conceptual Design Methods of High Mach Number UAVs in Near Space", the climb angle q and the unit residual power P S The coupling relationship can be expressed as:

[0166]

[0167] According to equations (28) and (29), we can obtain:

[0168]

[0169] The energy gain per unit mass of fuel consumed during the climb of a combined-power vehicle is defined as the energy conversion efficiency η, which can be expressed by the formula:

[0170]

[0171] Energy level is determined by He i Change to He i+1 When, He i For the i-th energy level, He i+1 The energy level is the (i+1)th level; the corresponding mass change is the fuel consumption mass Δm at this stage. f It can be obtained by integrating the following formula:

[0172]

[0173] According to formula (6) in step 2, the relationship between the aircraft thrust F and specific impulse can be expressed as follows:

[0174]

[0175] Based on the literature "Modeling and Multi-System Correlation Characteristics Analysis of Aerospace Vehicles Oriented to Control Integration", formula (9) in step 2 can be refined, and the corresponding specific impulse expression is:

[0176] I sp =60.09Ma 2 -1225Ma+7568;

[0177] Substituting equations (31), (33), and (29) into equation (32), we get:

[0178]

[0179] The thrust-to-weight ratio parameter is obtained from step 3-2. Substituting into formula (34), the energy height can be calculated from He. i Change to He i+1 When the mass change is equal to the mass of fuel consumed during that stage, Δm f ;

[0180] Assuming that the fuel consumed by the combined-powered vehicle from horizontal takeoff to the start of hypersonic climb accounts for 5% of the total takeoff mass of the combined-powered vehicle, therefore W in equation (34) i The initial value W0 = 0.95 * W TO The fuel consumption mass between each pair of adjacent isoenergy altitude lines can be iteratively calculated. Summing up the fuel consumption mass for each segment yields the total fuel consumption for hypersonic ascent, i.e., SDm. f (i)=W F Therefore, the relationship between the real-time changing fuel mass of the launch vehicle and the speed v and altitude h of the launch vehicle can be obtained.

[0181] Step 4: Based on the real-time changing fuel mass of the launch vehicle and its flight speed v and altitude h, discretize the climb trajectory for speed and altitude to obtain several sets of one-to-one corresponding data on the launch vehicle's flight speed v, altitude h, and fuel mass W. F Relationship;

[0182] Therefore, the total takeoff mass W for each set of relationships can be calculated according to formula (2). TO For example, in this embodiment, it is assumed that the structural mass ratio is... Choose 0.2, load mass W P =20t, machine weight WC Ignoreable, obtain The total takeoff mass W can then be calculated. TO ;

[0183] The total takeoff mass W obtained TO In the middle, select W TO The minimum velocity v and height h are taken as the optimal separation point.

[0184] Example 2:

[0185] Based on the separation point optimization method provided in Example 1, simulation calculations were performed on the separation speed and separation height at the separation point:

[0186] The simulation calculation aims to send a 20t mission-level aircraft into a parking orbit (at an altitude of 200km) using a launch vehicle. According to the overall design parameters of the combined propulsion launch vehicle consisting of the launch vehicle and the mission vehicle, the service ceiling of the launch vehicle is approximately 35km. The basic parameters of the launch vehicle are shown in Table 1.

[0187] Table 1

[0188]

[0189] Regarding energy altitude, when the launch vehicle's speed is relatively high, altitude has a relatively small impact on fuel consumption during the orbital insertion phase. (Velocity to remaining mass ratio m) F and the total mass m of the combined propulsion vehicle TO Relationship such as Figure 7 As shown, where m TO =W TO / g;μ=m TO g / W TO ;

[0190] Combining formula (34) and step 4 of Example 1, several sets of one-to-one corresponding flight speeds v, altitudes h, and fuel masses W of the launch vehicle are obtained from a given altitude condition. F The relationship between velocity v and altitude h forms an interstage separation optimization window. Within this window, the total takeoff mass W is found. TO The minimum value of the velocity v and altitude h corresponding to this minimum value is the optimal separation point. In this embodiment, the parameters of the separation point are V = 1300 m / s and h = 26 km. At this time, V = 1300 m / s is the separation velocity, and h = 26 km is the separation altitude. The total takeoff mass of the combined-powered launch vehicle is 216.3 tons, of which the takeoff mass of the launch vehicle is 158.7 tons and the takeoff mass of the mission vehicle is 57.5 tons. Figure 8 As shown.

[0191] Example 3:

[0192] The sensitivity of the overall design parameters of the combined-powered launch vehicle corresponding to the separation point obtained from the simulation calculation in Example 2 was verified:

[0193] The overall design parameters include: the structural mass ratio m of the launch vehicle. E Specific impulse of launch vehicles and specific impulse of mission vehicles;

[0194] (1) Structural mass ratio of launch vehicle m E It is one of the conceptual design parameters for combined-powered launch vehicles, and is related to the structural mass ratio (m) of the launch vehicle. E By applying bias, the effect on the interstage separation output parameters can be obtained, while the types of interstage separation output parameters remain unchanged, such as... Figure 9 As shown.

[0195] from Figure 9 As can be seen from this, when the structural mass of the launch vehicle is greater than m E As the structural mass ratio increases from 80% to 120%, the total takeoff mass of the combined-powered vehicle shows a significant increasing trend; for example, when the structural mass ratio is 80%, the total takeoff mass is 189 tons; while when the structural mass ratio reaches 120%, the total takeoff mass increases to 243.6 tons. This indicates that the increase in structural mass leads to an increase in the overall mass of the combined-powered vehicle. This is because a higher structural mass ratio means that more materials and components are needed to support the structure of the combined-powered vehicle, and more fuel is needed to propel the combined-powered vehicle to meet design requirements. For the simulation object obtained in Example 2, the rate of change is approximately 13%.

[0196] (2) Specific impulse is one of the metrics for evaluating the efficiency of a propulsion system; increasing specific impulse can reduce the amount of fuel required, thereby increasing the payload or improving the climb and transport capabilities of a combined-powered launch vehicle. Using the specific impulse of the launch vehicle as a design parameter, adjusting it can reveal its impact on interstage separation output parameters, such as... Figure 10 and Figure 11 As shown;

[0197] from Figure 10 It can be seen from this that when the specific impulse I of the mission-level aircraft... SP2 As the specific impulse changes from 80% to 120%, both the optimal separation altitude and velocity will change. The higher the specific impulse of the mission-grade aircraft, the lower the separation velocity and the lower the separation altitude; the total takeoff mass changes from 229.4t to 205.4t, a change rate of 5%.

[0198] from Figure 11 It can be seen from this that the specific impulse I of the launch vehicle is... SP1 (i.e. I SPThe impact of changes in the ramjet mode specific impulse on the combined-powered launch vehicle's performance parameters is analyzed. When the specific impulse of the launch vehicle's ramjet mode is 80%, the total takeoff mass of the combined-powered launch vehicle is 222.5 tons. As the specific impulse of the launch vehicle's ramjet mode gradually increases to 120%, the total takeoff mass gradually decreases. The optimal separation speed and altitude also increase accordingly. For example, when the specific impulse increases to 120%, the total takeoff mass decreases to 211.8 tons. This indicates that increasing the specific impulse of the launch vehicle's ramjet mode can improve engine efficiency, thereby reducing the required fuel and leading to a decrease in the total takeoff mass. Therefore, by conducting a sensitivity analysis of the impact of changes in the specific impulse of the launch vehicle's ramjet mode on the performance parameters of the combined-powered launch vehicle, a better understanding of the key factors in the design of combined-powered launch vehicles can be achieved. This analysis helps optimize the vehicle during the design phase to meet specific performance requirements and mission objectives.

[0199] For ease of description, spatial relative terms such as "above," "on top of," "on the upper surface of," "above," etc., are used herein to describe the spatial positional relationship of a device or feature as shown in the figures to other devices or features. It should be understood that spatial relative terms are intended to encompass different orientations in use or operation beyond the orientation of the device as described in the figures. For example, if the device in the figures were inverted, a device described as "above" or "on top of" other devices or structures would subsequently be positioned as "below" or "under" other devices or structures. Thus, the exemplary term "above" can include both "above" and "below." The device may also be positioned in other different ways (rotated 90 degrees or in other orientations), and the spatial relative descriptions used herein will be interpreted accordingly.

[0200] Furthermore, it should be noted that the use of terms such as "first" and "second" to define components is merely for the purpose of distinguishing the corresponding components. Unless otherwise stated, the above terms have no special meaning and therefore should not be construed as limiting the scope of protection of this invention.

[0201] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. An optimization method for the two-stage-to-orbit separation point of a combined-powered launch vehicle based on energy climb, wherein the combined-powered launch vehicle includes a launch-stage vehicle and a mission-stage vehicle; the separation point includes separation velocity and separation altitude; Its features are, The specific steps of this separation point optimization method are as follows: Step 1: Perform mass modeling on the combined-powered launch vehicle to determine the selection of the separation point and the total takeoff mass W of the combined-powered launch vehicle. TO Relationship: The selected separation point ensures that the total takeoff mass W TO Minimum; Step 2: Model the propulsion system of the launch vehicle and establish its specific impulse I. SP Fuel consumption rate / fuel consumption Relationship; Step 3: Before the mission-class aircraft separates from the launch vehicle, calculate the relationship between the real-time changing fuel weight of the launch vehicle and the speed v and altitude h of the launch vehicle based on the energy altitude. Step 4: Based on the relationship between the real-time changing fuel weight of the launch vehicle and its flight speed v and altitude h, discretize the climb trajectory for speed and altitude to obtain several sets of one-to-one corresponding data on the launch vehicle's flight speed v, altitude h, and fuel weight W. F The relationship between the two sets of relationships was calculated, and the total takeoff mass W for each set of relationships was obtained. TO In all the total takeoff mass W obtained TO In the middle, select W TO The minimum velocity v and height h are taken as the optimal separation point.

2. The method for optimizing the two-stage orbital separation point of a combined-powered launch vehicle based on energy climbing, as described in claim 1, is characterized in that... In step 1, the total takeoff mass W of the combined-powered launch vehicle TO Including the structural weight W of the launch vehicle E The fuel weight W carried by the launch vehicle F The maintenance weight W of the launch vehicle C Payload weight W of the launch vehicle P : IN TO =In E +W F +W P +W C (1) Among them, the payload weight W of the launch vehicle P This refers to the total weight of the mission-class aircraft; the payload weight W of the launch vehicle-class aircraft. P Maintenance weight W of launch vehicle C For the given design requirements, under the given W P and W C Under the condition of; adopting structural mass ratio ratio of fuel mass Total takeoff mass W TO Represented as: Therefore, the total takeoff mass W of the combined-powered launch vehicle TO The estimation needs to be combined with the structural mass ratio fuel mass ratio Combined-fuel launch vehicles use the separation point as the mission interface; differences in the separation point affect the fuel weight W of the launch vehicle. F Therefore, the principle for selecting the separation point is: the selected separation point should ensure that the total takeoff mass W... TO Minimum.

3. The method for optimizing the two-stage orbital separation point of a combined-powered launch vehicle based on energy climbing, as described in claim 2, is characterized in that... The value range of the structural mass ratio is 0.2 to 0.

75.

4. The method for optimizing the two-stage orbital separation point of a combined-powered launch vehicle based on energy climbing, as described in claim 1, is characterized in that... In step 2, the specific impulse I of the launch vehicle SP Fuel consumption rate / fuel consumption The relationship is: Where F represents the thrust generated by the propulsion system of the launch vehicle; g represents the gravitational acceleration at sea level.

5. The method for optimizing the two-stage orbital separation point of a combined-powered launch vehicle based on energy climbing, as described in claim 4, is characterized in that... Specific impulse I of launch vehicle SP Represented as: in, These are five different coefficients: M is the flight Mach number of the launch vehicle, and ∝ is the angle of attack of the launch vehicle.

6. The method for optimizing the two-stage orbital separation point of a combined-powered launch vehicle based on energy climbing, as described in claim 1, is characterized in that... The specific steps for step 3 are as follows: Step 3-1: Perform a force analysis on the combined-powered launch vehicle. Let the angle of attack α be the angle between the incoming flow velocity and the chord line of the launch vehicle, the climb angle be θ, the thrust F be forward along the chord line, the lift L be perpendicular to the incoming flow direction, the drag D be parallel to the incoming flow direction, the total weight W of the combined-powered launch vehicle be vertically downward, the acceleration of the combined-powered launch vehicle be a, the velocity be V, and the altitude be h. This yields the thrust-to-weight ratio parameter. Step 3-2, calculate the acceleration 'a' of the launch vehicle during the climb phase: Where V1 is the velocity at the intersection of the previous level isoenergy altitude line and the climbing curve, and V2 is the velocity at the intersection of the next level isoenergy altitude line and the climbing curve; h1 is the altitude at the intersection of the previous level isoenergy altitude line and the climbing curve, and h2 is the altitude at the intersection of the next level isoenergy altitude line and the climbing curve. Substituting formula (23) into formula (18), the thrust-to-weight ratio parameter can be calculated. Step 3-3: Calculate the fuel consumption mass Δm of the launch vehicle during the climb phase. f : Wherein, energy conversion efficiency η; He is the energy level. He i For the i-th energy level, He i+1 The energy height of level i+1; Fuel consumption rate / fuel consumption; P S Residual power per unit weight; I SP Specific impulse for launch vehicles; The thrust-to-weight ratio parameter is obtained from step 3-2. Substituting into formula (34), the energy height can be calculated from He. i Change to He i+1 When the mass change is equal to the mass of fuel consumed during that stage, Δm f ; The fuel consumption mass between each two adjacent isoenergy altitude lines can be iteratively calculated. Summing up the fuel consumption mass for each segment yields the total fuel consumption for hypersonic climb, i.e., SDm. f (i)=W F This allows us to obtain the relationship between the real-time changing fuel weight of the launch vehicle and its flight speed v and altitude h.

7. The method for optimizing the two-stage orbital separation point of a combined-powered launch vehicle based on energy climbing, as described in claim 6, is characterized in that... The specific process of step 3-1 is as follows: According to Newton's second law, force analysis of the combined-powered vehicle along the body direction yields the following results: Force analysis of the combined propulsion vehicle in the direction perpendicular to the body: L=Fsinq+Wcosq=Fsina+Wcosq (14) Combining equations (13) and (14), we can obtain: Since α≈θ, formula (15) can be transformed into: cosa≈1, sinα≈0, formula (16) can be transformed into: Further simplification yields:

8. The method for optimizing the two-stage orbital separation point of a combined-powered launch vehicle based on energy climbing, as described in claim 6, is characterized in that... The specific process of step 3-2 is as follows: The displacement component of the climb segment along the body direction can be expressed as: By deriving formula (19), we can obtain: Further simplification yields: Since the displacement component of the climb segment parallel to the body direction can also be expressed as: Therefore, the acceleration 'a' during the ascent phase can be expressed as:

9. The method for optimizing the two-stage orbital separation point of a combined-powered launch vehicle based on energy climbing, as described in claim 7, is characterized in that... The specific process of step 3-3 is as follows: Multiplying both sides of equation (13) by the velocity V, we get: Wherein, the vertical component of velocity V is the climb velocity, which can be expressed as: Combining equations (24) and (25), we can obtain: In equation (26), the right side of the equation represents the residual power per unit weight, P. S , can be represented as follows: Combining equations (26) and (27), we can obtain: Due to the climb angle q and the unit residual power P S The coupling relationship can be expressed as: According to equations (28) and (29), we can obtain: The energy gain per unit mass of fuel consumed during the climb of a combined-power vehicle is defined as the energy conversion efficiency η, which can be expressed by the formula: Energy level is determined by He i Change to He i+1 When the mass change is equal to the mass of fuel consumed during that stage, Δm f It can be obtained by integrating the following formula: The relationship between aircraft thrust F and specific impulse can be expressed as: Substituting equations (31), (33), and (29) into equation (32) yields equation (34).

10. The method for optimizing the two-stage orbital separation point of a combined-powered launch vehicle based on energy climbing, as described in claim 6, is characterized in that... Let the fuel consumed by the combined-powered vehicle from horizontal takeoff to the start of hypersonic climb account for 5% of the total takeoff mass of the combined-powered vehicle, therefore W in equation (34) i The initial value W0 = 0.95 * W TO .

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