Layout design method for long-term on-orbit large-volume ultra-low-orbit aircraft

Abstract

A layout design method for a long-term on-orbit large-volume ultralow-orbit aircraft belongs to the technical field of aerodynamic design of aerospace aircrafts, and comprises the following steps: designing a large-volume fuselage with a large slenderness ratio; designing a capture section molded surface of the air inlet duct; a compression section molded surface of the air inlet channel is designed; the machine body and the air inlet channel are integrated; arranging and fixing solar cell wings; and arranging the folded solar cell wing. The fuselage with the approximately-circular section can cooperate with an air suction type electric propulsion power system in a pneumatic assisting mode to achieve attitude adjustment and orbit maneuvering of the aircraft. The invention aims to develop an ultra-low orbit aircraft capable of residing for a long time, and develops a layout design method of the aircraft capable of residing for a long time on an ultra-low orbit, a large-volume fuselage can ensure sufficient load carrying space, multi-task requirements of the ultra-low orbit aircraft are met, and meanwhile, pneumatic assistance can be provided for in-orbit operation of the aircraft.

Description

Technical Field

[0001] This invention relates to a layout design method for a large-volume ultra-low orbit vehicle that can operate in orbit for a long time, and belongs to the field of aerospace vehicle aerodynamic design technology. Background Technology

[0002] Ultra-low Earth orbit (ULE) (typically referring to the airspace 100-300 kilometers above the Earth) is becoming a focus of global aerospace attention due to its unique space environment and potential applications. This orbit, significantly lower than traditional medium and low Earth orbit satellites, offers unprecedented advantages: First, the reduced orbital altitude greatly improves the resolution of optical observations, enabling satellites to achieve higher-precision imaging capabilities without increasing size or cost, significantly enhancing the ability to capture detailed Earth observations; second, the shorter distance between satellite and Earth reduces communication latency and signal loss, ensuring real-time communication and high-speed data transmission; furthermore, the significantly reduced size and weight of satellites significantly lowers the manufacturing and launch costs of individual satellites, making it possible to build low-cost, high-density satellite constellations; finally, the thin atmosphere and low radiation dose in this orbital airspace are conducive to high-precision geophysical measurements and provide convenient deorbiting capabilities, reducing space debris.

[0003] Despite the immense potential of ultra-low Earth orbit (ULO) vehicles, the complex and variable atmospheric environment of their airspace presents significant challenges to their design and application. The atmospheric density of this orbit (10⁻⁶) is... -7 ~10 -12 kg / m 3 Although the atmospheric drag is far below that of traditional airspace (making it negligible relative to aircraft), it is significantly higher than that of low and medium Earth orbits. This thin atmosphere generates significant aerodynamic drag, and coupled with the nonlinear and uncertain characteristics of the atmospheric environment—where atmospheric density varies not only with orbital altitude but also can differ considerably even at the same altitude—spacecraft must continuously overcome drag and make attitude and orbital corrections. Traditional spacecraft, limited by finite propellant, cannot sustain long-term operation in this environment. Therefore, developing spacecraft capable of stable long-term operation in ultra-low Earth orbit is crucial to realizing the advantages of this orbit.

[0004] To enable ultra-low orbit (ULE) vehicles to remain in orbit for extended periods and perform a variety of on-orbit missions, and to fully leverage the advantages of ULE orbit, it is imperative to propose a layout design method for large-volume ULE vehicles capable of long-term on-orbit operation. Summary of the Invention

[0005] The purpose of this invention is to develop a long-term stationary ultra-low orbit (ULO) vehicle and to provide a layout design method for a large-volume ULE vehicle capable of long-term on-orbit operation. This layout primarily employs a large-volume aircraft-like fuselage, with the fuselage and air-breathing electric propulsion system highly integrated, resulting in a layout design method for a vehicle capable of long-term stationary operation in ULE. In this layout, the large-volume fuselage ensures sufficient payload space to support the multi-mission requirements of the ULE vehicle, while also providing aerodynamic assistance for on-orbit operation. A brief overview of the invention is provided below to offer a basic understanding of certain aspects of the invention. It should be understood that this overview is not an exhaustive summary of the invention. It is not intended to identify key or essential parts of the invention, nor is it intended to limit the scope of the invention.

[0006] The technical solution of the present invention:

[0007] A layout design method for a large-volume ultra-low Earth orbit vehicle with long-term on-orbit capability includes the following steps:

[0008] Step 1, Design a large-capacity fuselage: Under the constraint of meeting the predetermined volume requirements, determine the aspect ratio of the main fuselage and adopt a near-circular cross-section configuration to make the fuselage surface smoothly transition to reduce wind resistance;

[0009] Step 2, design the capture section profile of the air intake: An air-breathing electric propulsion system is installed under the fuselage. A capture section is set at the front end of the air intake within the air-breathing electric propulsion system. The capture section is a cylindrical structure with a honeycomb grid inside. The length of the capture section and the size of the honeycomb grid are determined according to the overall size of the aircraft and the atmospheric capture flow requirements.

[0010] Step 3, design the compression section profile of the air intake: The compression section is connected behind the capture section. The compression section adopts a parabolic profile, and its length-to-length ratio and contraction ratio are selected according to the propulsion requirements of the aircraft.

[0011] Step 4, integrate the fuselage and air intake: integrate the air intake and fuselage structures to ensure that the fuselage volume is not affected and maintain the compactness of the aircraft configuration;

[0012] Step 5, Arrange fixed solar cell wings: Fixed solar cell wings are symmetrically and fixedly connected on both sides of the fuselage. The fixed solar cell wings are swept back along the flight direction, and the sweep angle θ1 ranges from 0° to 80°.

[0013] Step 6, Arrange foldable solar cell wings: Foldable solar cell wings are symmetrically arranged on the outside of the fixed solar cell wings. The foldable solar cell wings adopt a swept-back configuration, and their swept-back angle θ2 satisfies 0°≤θ2≤θ1.

[0014] This invention integrates the layout features of large-volume aircraft and spacecraft. This layout can achieve on-orbit aerodynamic assistance through aerodynamic configuration, and at the same time, it combines an air-breathing electric propulsion system to cope with the drag of the thin atmosphere, so as to achieve long-term operation in ultra-low orbit. It aims to provide a solution reference for the layout design and development direction of ultra-low orbit aircraft.

[0015] Preferably, the foldable solar cell wing is mounted on the fixed solar cell wing via a folding mechanism.

[0016] Preferably, the honeycomb grid has a regular hexagonal structure.

[0017] Preferably, the inlet diameter of the compression section is the same as the diameter of the capture section.

[0018] Preferably, in steps 5 and 6, the dimensional parameters of the fixed solar cell wing and the folding solar cell wing are designed to match the overall energy requirements and structural layout of the aircraft.

[0019] Preferably, it also includes arranging solar cells on the surface of the rear profile of the fuselage.

[0020] The present invention has the following beneficial effects:

[0021] 1. Aerodynamic Advantages: The ultra-low orbit vehicle layout provided by this invention, with its near-circular cross-section and large-volume fuselage, offers significant aerodynamic advantages. The near-circular cross-section fuselage configuration is more effective at mitigating drag in the thin atmosphere and can provide aerodynamic assistance for the vehicle's on-orbit operation, supporting attitude adjustment and orbital transfer.

[0022] 2. Structural Advantages: The ultra-low orbit vehicle layout provided by this invention utilizes a near-circular cross-section, large-volume fuselage with greater loading space, meeting the requirements for carrying various mission payloads. Simultaneously, the near-circular cross-section fuselage provides a larger surface area, ensuring the placement of solar cells and high-power solar energy absorption and conversion. The swept-back layout of fixed and folding solar panels helps reduce wave drag, thereby reducing atmospheric drag.

[0023] 3. Performance Advantages: The ultra-low orbit (ULO) vehicle layout provided by this invention features a highly integrated fuselage and air-breathing electric propulsion system, enabling the vehicle to operate in orbit for extended periods and offering significant performance advantages. The working propellant for the air-breathing electric propulsion system can be self-supplied within the ULE airspace, providing a stable guarantee for the long-term on-orbit operation of the ULE vehicle.

[0024] 4. This invention patent has the significant features of simple overall structure, sufficient loading space and low economic cost, providing a reliable aircraft layout for the exploration of ultra-low orbit airspace. Attached Figure Description

[0025] Figure 1 This is a perspective view of the aircraft described in the specific embodiment;

[0026] Figure 2 This is a perspective view of the fuselage described in the specific embodiment;

[0027] Figure 3 This is a perspective view of the air-breathing electric propulsion system described in the specific embodiment;

[0028] Figure 4 This is a schematic diagram of the honeycomb grid structure described in the specific embodiment;

[0029] Figure 5 This is a schematic diagram of the structure when the honeycomb grid described in the specific embodiment is a regular hexagonal structure;

[0030] Figure 6 This is a perspective view of the compression section described in the specific implementation method;

[0031] Figure 7 This is a schematic diagram of the structure of the fixed solar cell wing described in the specific embodiment;

[0032] Figure 8 This is a schematic diagram of the structure of the folding solar cell wing described in the specific embodiment;

[0033] Figure 9 This is a top view of the aircraft described in the specific embodiment;

[0034] Figure 10 This is a front view of the aircraft described in the specific embodiment;

[0035] Figure 11 This is a rear view of the aircraft described in the specific embodiment;

[0036] Figure 12 This is a side view of the aircraft described in the specific embodiment.

[0037] In the diagram, 1-fuselage, 2-air-breathing electric propulsion system, 3-fixed solar cell wing, 4-folding solar cell wing, 5-folding mechanism, 6-fuselage back profile, 7-fuselage belly profile, 8-capture section, 9-compression section, 10-honeycomb grid. Detailed Implementation

[0038] To make the objectives, technical solutions, and advantages of this invention clearer, the invention is described below with reference to specific embodiments shown in the accompanying drawings. However, it should be understood that these descriptions are merely exemplary and not intended to limit the scope of the invention. Furthermore, descriptions of well-known structures and technologies are omitted in the following description to avoid unnecessarily obscuring the concept of the invention.

[0039] The connections mentioned in this invention are divided into fixed connections and detachable connections. Fixed connections (i.e., non-detachable connections) include, but are not limited to, conventional fixed connection methods such as folded connections, riveted connections, adhesive connections, and welded connections. Detachable connections include, but are not limited to, conventional disassembly methods such as threaded connections, snap-fit ​​connections, pin connections, and hinged connections. When a specific connection method is not explicitly defined, it is assumed that at least one existing connection method can always be found to achieve the function, and those skilled in the art can choose according to their needs. For example, a welded connection can be chosen for fixed connections, and a hinged connection can be chosen for detachable connections.

[0040] Combination Figures 1-12 This embodiment describes a layout design method for a long-term on-orbit, large-volume ultra-low orbit (ULO) spacecraft. The spacecraft includes a fuselage 1, an air-breathing electric propulsion system 2, fixed solar panels 3, folding solar panels 4, and a folding mechanism 5. Fixed solar panels 3 are mounted on the left and right sides of the fuselage 1. Folding solar panels 4 are mounted on the fixed solar panels 3 via the folding mechanism 5. The air-breathing electric propulsion system 2 is located below the fuselage 1. An air intake is provided within the air-breathing electric propulsion system 2, and a capture section 8 and a compression section 9 are sequentially arranged within the air intake.

[0041] The method includes the following steps:

[0042] Step 1, Design a large-volume fuselage 1: Under the constraint of meeting the predetermined volume requirements, appropriately increase the aspect ratio of fuselage 1 to reduce the aircraft's frontal area and thus reduce frontal drag. Without considering the solar array, the aspect ratio of a conventional satellite is approximately 0.8~1.5. In this embodiment, the aspect ratio of fuselage 1 is selected in the range of 4~6, and a near-circular cross-section configuration is adopted to make the surface lines of fuselage 1 smooth, thereby reducing the aircraft's frontal and lateral drag. The aspect ratio of fuselage 1 without considering the solar array is not less than 4; in this embodiment, the aspect ratio of fuselage 1 is selected as 6.

[0043] Specifically, based on the head payload volume requirements, the dorsal fuselage surface 6 and the ventral fuselage surface 7 are designed respectively. Based on the dorsal fuselage surface 6 and the ventral fuselage surface 7, and according to the overall volume requirements of the ultra-low orbit vehicle, a large-volume fuselage 1 with an approximately circular cross-section is designed. This near-circular cross-section not only helps reduce aerodynamic drag but also facilitates subsequent integration with the air-breathing electric propulsion system 2.

[0044] Step 2, design the 8-face shape of the intake duct's capture section:

[0045] The air-breathing electric propulsion system 2 is installed below the fuselage 1. At the same time, in order to maximize energy acquisition, solar cells are arranged on the surface of the fuselage back surface 6. The large surface area provided by the near-circular fuselage ensures the absorption and conversion of solar energy at high power.

[0046] This embodiment only provides the layout design of the air intake. The pressurization storage device and electric thruster of the air-breathing electric propulsion power system 2 are only illustrated and do not involve specific configurations and types.

[0047] A capture section 8 is provided at the front end of the air intake in the air-breathing electric propulsion system 2. The capture section 8 is a cylindrical structure with a diameter of φ1, and a honeycomb grid 10 is arranged inside. The length of the capture section 8 and the size of the honeycomb grid 10 are determined according to the overall size of the aircraft and the atmospheric capture flow requirements. This structure reduces the outward reflection and escape of captured atmospheric molecules, thereby enhancing the capture efficiency of thin atmosphere. In this embodiment, as shown... Figure 5 As shown, the honeycomb grid 10 is preferably a regular hexagonal structure, with the inscribed circle diameter of a single grid being φ2=0.067φ1, and the length of the capturing segment 8 being 3.3φ1. In specific engineering implementation schemes, the optimal design needs to be carried out according to the specific scheme. This embodiment is only for configuration illustration.

[0048] Step 3: Design the profile of the compression section 9 of the inlet: The compression section 9 is connected behind the capture section 8. The compression section 9 adopts a parabolic profile to compress and pressurize the atmospheric molecules collected by the capture section 8. The control parameters of the parabolic profile are selected during the optimization design process and need to be determined in conjunction with key performance parameters such as the compression ratio and aspect ratio of the compression section. The inlet diameter of the compression section 9 is consistent with the diameter φ1 of the capture section 8, that is, the inlet diameter φ3 of the compression section 9 of the inlet is φ1. The aspect ratio and contraction ratio of the compression section profile need to be selected and designed in conjunction with the pressure and flow requirements of the working fluid in the propulsion system of the aircraft.

[0049] Step 4, integrate fuselage 1 and air intake: In order to ensure the compactness of the aircraft configuration and the smoothness of the aerodynamic shape, the air intake and fuselage 1 are designed as an integrated unit to ensure that the volume of fuselage 1 is not affected and to maintain the compactness of the aircraft configuration.

[0050] Specifically, the upper side of the air intake of the air-breathing electric propulsion system 2 is matched with the fuselage belly surface 7 described in step 1, and the side of the air intake smoothly transitions with the side surfaces of the fuselage 1. This integrated design ensures that the internal volume of the fuselage 1 is not affected to support the mission payload, while achieving deep integration of the propulsion system and the main structure.

[0051] Step 5, Arrange fixed solar panels 3: Fixed solar panels 3 are symmetrically and fixedly connected on both sides of the fuselage 1. The fixed solar panels 3 are swept back along the flight direction, and the sweep angle θ1 ranges from 0° to 80°. When the sweep angle is greater than 80°, the usable fixed wing area is too small, and its economy and effectiveness cannot be guaranteed. In the specific scheme, the specific value needs to be determined in combination with the solar panel area requirements and structural strength requirements. The folding solar panel 4 is subsequently installed on the fixed solar panel 3 through the folding mechanism 5. The specific dimensions of the fixed solar panel 3, such as the root length L1, wingtip length, span width W1, and thickness H1, need to be matched and designed in combination with the overall energy requirements and structural layout of the aircraft.

[0052] Step 6, Arrange the foldable solar panel wing 4: Foldable solar panels wing 4 are symmetrically arranged on the outside of the fixed solar panel wing 3. The foldable solar panel wing 4 also adopts a swept configuration to reduce wave drag, and its sweep angle θ2 satisfies 0°≤θ2≤θ1. The root length of the foldable solar panel wing 4 is L2, the wingtip length is L3, the span is W2, and the thickness is H2. To reduce the envelope space of the aircraft during the launch phase and reduce the launch difficulty, a dedicated folding mechanism 5 is provided. The dimensional parameters of the fixed solar panel wing 3 and the foldable solar panel wing 4 are ultimately designed to match the overall energy requirements and structural layout of the aircraft.

[0053] This embodiment describes a layout design method for an ultra-low orbit (ULO) aircraft with a near-circular cross-section fuselage with a large aspect ratio, swept solar panels, and an air-breathing electric propulsion system. The aim is to reduce the aircraft's frontal and lateral drag in the thin atmosphere of the ULE airspace by using a near-circular cross-section fuselage with a large aspect ratio; the swept solar panels reduce the probability and force of head-on collisions with atmospheric molecules to a certain extent; at the same time, the near-circular cross-section fuselage and swept solar panels can obtain the aerodynamic characteristics of the thin atmosphere by adjusting the angle of attack. This aerodynamic force can work together with the thrust generated by the air-breathing electric propulsion system to assist the aircraft in slow attitude adjustment and orbital transfer.

[0054] This implementation successfully integrates the features of a large-volume aircraft fuselage and a spacecraft layout through the steps described above, and achieves a high degree of integration with the air-breathing electric propulsion system. This layout not only provides ample loading space and an efficient energy system, but also significantly enhances the long-term loiter capability in ultra-low Earth orbit environments through aerodynamic optimization design, providing a reliable platform for missions such as high-precision Earth observation and low-latency communication.

[0055] It should be noted that in the above embodiments, as long as the technical solutions are not contradictory, they can be permuted and combined. Those skilled in the art can exhaust all possibilities based on the mathematical knowledge of permutation and combination. Therefore, the present invention will not describe the technical solutions after permutation and combination one by one, but it should be understood that the technical solutions after permutation and combination have been disclosed by the present invention.

[0056] The above description is merely a preferred embodiment of the present invention and is not intended to limit the 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. A long-term on-orbit large-volume ultra-low-orbit vehicle layout design method, characterized by, It comprises the following steps: Step 1, design a large volume fuselage (1): under the constraint of meeting the predetermined large volume demand, determine the slenderness ratio of the fuselage (1), and adopt a near-circular cross-section configuration; Step 2, design the capture section (8) profile of the inlet: the lower part of the fuselage (1) is provided with an air-breathing electric propulsion power system (2), and the front end of the inlet in the air-breathing electric propulsion power system (2) is provided with a capture section (8), the capture section (8) is a cylindrical structure, and a honeycomb grid (10) is arranged inside the capture section (8), the length of the capture section (8) and the size of the honeycomb grid (10) are determined according to the overall size of the aircraft and the atmospheric capture flow demand; Step 3, design the compression section (9) profile of the inlet: connect the compression section (9) behind the capture section (8), the compression section (9) adopts a parabolic profile, and the length-to-width ratio and the contraction ratio are selected according to the propulsion demand of the aircraft; Step 4, integrated integration of fuselage (1) and inlet: the structure of the inlet and the fuselage (1) is designed integrally to ensure that the volume of the fuselage (1) is not affected and the compactness of the aircraft configuration is maintained; Step 5, arrange the fixed solar cell wing (3): symmetrically arrange the fixed solar cell wing (3) fixedly connected on both sides of the fuselage (1), the fixed solar cell wing (3) is in a backward-swept configuration along the flight direction, and the backward-swept angle θ1 is in the range of 0° to 80°; Step 6, arrange the folding solar cell wing (4): symmetrically arrange the folding solar cell wing (4) on the outside of the fixed solar cell wing (3), the folding solar cell wing (4) adopts a backward-swept configuration, and the backward-swept angle θ2 satisfies 0°≤θ2≤θ1.

2. The long-term on-orbit large-volume ultra-low-orbit vehicle layout design method according to claim 1, characterized in that: The folding solar cell wing (4) is installed on the fixed solar cell wing (3) through the folding mechanism (5).

3. The long-term on-orbit large-volume ultra-low-orbit vehicle layout design method according to claim 1, characterized in that: The honeycomb grid (10) is a regular hexagonal structure.

4. The long-term on-orbit large-volume ultra-low-orbit vehicle layout design method according to claim 1, characterized in that: The inlet diameter of the compression section (9) is consistent with the diameter of the capture section (8).

5. The long-term on-orbit large-volume ultra-low-orbit vehicle layout design method according to claim 1, characterized in that: In steps 5 and 6, the size parameters of the fixed solar cell wing (3) and the folding solar cell wing (4) are matched and designed according to the overall energy demand of the aircraft and the structure layout.

6. The long-term on-orbit large-volume ultra-low-orbit vehicle layout design method according to claim 1, characterized in that: It also includes arranging solar cell pieces on the surface of the fuselage back profile (6).

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