Yaw maneuvering sun alignment method for inclined low-orbit aircraft
By designing a yaw maneuver to the Japanese method in an inclined low-orbit vehicle, calculating the sun's angle and setting the aircraft attitude maneuver and windsurfing rotation, the shortcomings of the windsurfing method in the prior art inclined low-orbit vehicle are solved, and efficient windsurfing to the Japanese and energy supplements are achieved.
Patent Information
- Application Number
- CN202510226322.6
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-02-27
- Publication Date
- 2025-06-10
AI Technical Summary
The prior art method of using one-dimensional rotating windsurfing to Japan in inclined low-orbit aircraft is not perfect, and it is difficult to ensure long-term replenishment of windsurfing to Japan and energy.
A yaw maneuvering method is proposed to calculate the sun's altitude angle, the sun's azimuth angle and the sun's pitch angle, and design the aircraft's yaw attitude maneuvering target attitude angle and windsurfing rotation method to meet the constraints at different sun altitude angles and realize windsurfing sun.
It realizes windsurfing to the sun through a small amount of attitude maneuvering in an inclined low-orbit aircraft, reducing fuel consumption and ensuring the efficiency and energy replenishment of solar windsurfing to the sun.
Smart Images
Figure CN120122708A_ABST
Abstract
Description
Technical Field
[0001] The invention relates to a yaw maneuvering method for tilting a low-orbit aircraft towards the sun, belonging to the technical field of autonomous sun-facing design of sailboards of on-orbit aircraft. Background Art
[0002] In the existing research on the method of aiming the sailboard at the sun of an inclined low-orbit spacecraft, the sailboard is usually aimed at the sun by using a fixed sailboard and the inertial orientation of the spacecraft, or by using a two-dimensional sailboard rotation. In the case of a one-dimensional rotating sailboard, the spacecraft is usually in a sun-synchronous orbit. The current method of using a one-dimensional sailboard to aim the sailboard at the sun for an inclined low-orbit spacecraft is still imperfect. Summary of the invention
[0003] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art, a yaw maneuver method for tilted low-orbit aircraft to face the sun is proposed, and based on different solar altitude angles, corresponding attitude maneuvering angles and sailboard rotation methods are designed to achieve long-term on-orbit sailboard facing the sun and ensure energy replenishment.
[0004] The technical solution of the present invention is:
[0005] A method for yaw maneuvering a tilted low-orbit aircraft towards the sun comprises the following steps:
[0006] Step 1: Calculate the sun vector in the inertial system based on the time and spacecraft orbit information;
[0007] Step 2: According to the solar vector, calculate the solar altitude angle, solar azimuth angle and solar pitch angle of the spacecraft orbit system; wherein the solar altitude angle is the angle between the solar vector of the spacecraft orbit system and the YOZ plane of the spacecraft orbit system, the solar azimuth angle is the angle between the solar vector of the spacecraft orbit system and the XOZ plane of the spacecraft orbit system, and the solar pitch angle is the angle between the solar vector of the spacecraft orbit system and the XOY plane of the spacecraft orbit system; wherein the origin of the spacecraft orbit system is the center of mass of the spacecraft, the OZ axis points to the center of the earth, the OX axis points to the orbital plane perpendicular to the OZ axis pointing to the flight direction of the spacecraft, and the OY axis is perpendicular to the OZ and OX axes according to the right-hand rule;
[0008] Step 3: Based on the constraints and the solar altitude angle, solar azimuth angle and solar pitch angle in the spacecraft orbit system obtained in step 2, design the target attitude angle of the spacecraft yaw attitude maneuver; after the spacecraft maneuvers to the target attitude angle, design the windsurfing board rotation speed.
[0009] Furthermore, the constraints of the aircraft yaw attitude maneuver are:
[0010] Constraint 1: The aircraft always keeps its head facing the sun;
[0011] Constraint 2: The optimal incident angle of the solar panel is less than or equal to 20 degrees;
[0012] Constraint 3: The aircraft should minimize attitude maneuvers per orbit to ensure the aircraft's TT&C requirements;
[0013] Constraint 4: Under an inclined orbit, the variation range of the solar altitude angle in one Earth year is [-70, 70].
[0014] Furthermore, to meet all the constraints, the target attitude angle for the aircraft's yaw attitude maneuver is designed as follows:
[0015] 1) When the solar altitude angle is [-90, -50], the target attitude angle is [0, 0, 90];
[0016] 2) When the solar altitude angle is (-50, -20] and the solar pitch angle is [-90, 90], the target attitude angle is [0, 0, 50];
[0017] 3) When the solar altitude angle is (-50, -20] and the solar pitch angle is [90, 180] or [-180, -90], the target attitude angle is [0, 0, 130];
[0018] 4) When the solar altitude angle is (-20, 20], the target attitude angle is [0, 0, 0];
[0019] 5) When the solar altitude angle is (20, 50] and the solar pitch angle is [90, 180] or [-180, -90], the target attitude angle is [0, 0, -130];
[0020] 6) When the solar altitude angle is (20, 50] and the solar pitch angle is [-90, 90], the target attitude angle is [0, 0, -50];
[0021] 7) When the solar altitude angle is [-90, -50), the target attitude angle is [0, 0, -90].
[0022] Furthermore, after the aircraft maneuvers to the target attitude angle, the rotational speed of the solar panel is designed. Through capture and tracking, the normal of the solar panel is made to follow the projection of the solar vector on the normal plane of the solar panel, that is, to ensure the minimum angle between the solar panel and the sun and the highest efficiency of the solar panel facing the sun.
[0023] Furthermore, after the aircraft maneuvers to the target attitude angle, the rotational speed of the solar panel is designed as follows:
[0024] 1) If abs(beta) ≥ 50, the normal of the solar panel coincides with the X-axis of the local coordinate system; where beta is the solar altitude angle;
[0025] 2) If 50 > abs(beta) > 20, then:
[0026] Calculate the deviation zeta between the current angle and the desired angle of the sailboard. The desired angle of the sailboard is the angle that satisfies the condition that the normal vector of the sailboard follows the projection of the solar vector on the normal plane of the sailboard.
[0027] zeta = mod2pi(mod2pi(-alpha + pi))
[0028] In the formula, alpha is the solar elevation angle.
[0029] If abs(zeta) > 5, then perform sailboard capture.
[0030] If abs(zeta) ≤ 5, then the rotational speed of the sailboard is sign(S ox ) * w and perform tracking; where w is the orbital angular velocity of the aircraft, and S ox is the component of the solar vector in the aircraft orbital system on the x-axis of the aircraft orbital system.
[0031] 3) If abs(beta) ≤ 20, then:
[0032] Calculate the deviation zeta between the current angle and the desired angle of the sailboard.
[0033] If abs(zeta) > 5, then perform sailboard capture.
[0034] If abs(zeta) ≤ 5, then the rotational speed of the sailboard is w and perform tracking.
[0035] Furthermore, according to the solar vector, calculate the solar altitude angle, solar azimuth angle, and solar elevation angle in the aircraft orbital system, specifically:
[0036] Transfer the solar vector S i in the inertial system to the aircraft orbital system:
[0037] S o = A b←i * S i
[0038] In the formula, A b←i is the attitude transformation matrix from the inertial coordinate system to the aircraft orbital system, and S o is the solar vector in the aircraft orbital system;
[0039] Then the solar altitude angle beta is:
[0040] beta = arcsin(S y / S)
[0041] The solar azimuth angle gama is:
[0042] gama = arctan(Sy / S x )
[0043] The solar elevation angle alpha is:
[0044] alpha = arctan(-S z / S x ).
[0045] Furthermore, according to the time and the vehicle orbit information, calculate the solar vector S in the inertial system i :
[0046]
[0047] where, Ω r is the right ascension of the ascending node in the inertial system, with a value of 0; u r is the latitude amplitude in the inertial system, and i r is the orbital inclination in the inertial system.
[0048] Furthermore, use the first-order terms to calculate the solar vector in the inertial system:
[0049] u r = ω r + f r
[0050] ω r = (282.937347 + 0.32256206T JC ) × (π / 180)
[0051] f r = M r + 2e r sin(M r ) + 1.25e r 2 sin(2M r )
[0052] M r = (357.5291 + 35999.0502888889T JC ) × (π / 180)
[0053] i r = (23.439291 - 0.01300417T JC ) × (π / 180)
[0054] In the formula, ω r is the argument of perigee, taking the first-order term, with the unit of rad; M r is the mean anomaly, modulo 2π, with the unit of rad; T JCis the Julian century number from the current time to J2000.0; f r is the true anomaly angle, in rad.
[0055] The advantages of the present invention compared with the prior art are:
[0056] (1) The yaw maneuver method for the inclined orbital vehicle provided by the present invention solves the problem of how to achieve solar alignment of the sailboard through a small amount of attitude maneuvers when the inclined orbital vehicle is only equipped with one-dimensional rotating solar wings.
[0057] (2) The yaw maneuver method for the inclined orbital aircraft provided by the present invention is based on different solar altitude angles. By designing the attitude maneuvering angle, it is ensured that the direction of the solar vector is as parallel to the normal surface of the sailboard as possible, while reducing the fuel consumption caused by the aircraft attitude maneuvering. At the same time, it is ensured that the sun protection surface of the aircraft head is facing the sun to prevent the aircraft tail thruster from being exposed to the sun.
[0058] (3) The method for the yaw maneuver of the inclined orbit vehicle towards the sun provided by the present invention is simple and easy to apply in engineering. BRIEF DESCRIPTION OF THE DRAWINGS
[0059] Various other advantages and benefits will become apparent to those of ordinary skill in the art by reading the detailed description of the preferred embodiments below. The accompanying drawings are only for the purpose of illustrating the preferred embodiments and are not to be considered as limiting the present invention. Moreover, the same reference symbols are used throughout the accompanying drawings to represent the same components. In the accompanying drawings:
[0060] Figure 1 The present invention is a flow chart of a method for performing a yaw maneuver toward the sun for a tilted low-orbit aircraft according to an embodiment of the present invention. DETAILED DESCRIPTION
[0061] The exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although the exemplary embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided to enable a more thorough understanding of the present disclosure and to fully convey the scope of the present disclosure to those skilled in the art.
[0062] The present invention proposes a method for tilting a low-orbit aircraft to perform a yaw maneuver to face the sun, which is used for an on-orbit aircraft to tilt the sailboard of the low-orbit aircraft to yaw to face the sun. Figure 1 As shown, the method specifically comprises the following steps:
[0063] Step 1: Calculate the sun vector S based on the time and spacecraft orbit information i
[0064] According to the performance constraints of the aircraft-borne computer, the design uses first-order quantities to calculate the solar vector. The semi-major axis a of the solar orbit parameters based on the J2000 geocentric inertial coordinate system (solar perspective) r (unit: m), eccentricity e r , orbital inclination i r (unit: rad), right ascension of the ascending node Ω r (unit: rad), argument of periapsis ω r (unit: rad), mean anomaly M r (unit: rad), and the Julian century number T from the current moment to J2000.0 JC are given by the following formulas:
[0065] a r = 1.00000102 (AU)
[0066] e r = 0.01670862 - 0.00004204T JC
[0067] i r = (23.439291 - 0.01300417T JC ) × (π / 180)
[0068] Ω = 0
[0069] ω r = (282.937347 + 0.32256206T JC ) × (π / 180)
[0070] M r = (357.5291 + 35999.0502888889T JC ) × (π / 180)
[0071] where AU is the astronomical unit length, with a value of 1.49597870 × 10^11 m. M r needs to be modulo 2π.
[0072] The true anomaly f r (unit: rad), and the latitude amplitude u r (unit: rad, i.e., the ecliptic longitude) are given by the following formulas:
[0073] f r = M r + 2e r sin(M r ) + 1.25e r 2 sin(2M r )
[0074] u r = ω r + f r
[0075] In the formula, u r needs to be modulo 2π.
[0076] Inertial system solar vector S i :
[0077]
[0078] Among them, Ω r is the right ascension of the ascending node, u r is the latitude amplitude, i r is the orbital inclination, and i represents the inertial system.
[0079] Step 2: Calculate the solar altitude angle beta, solar azimuth angle gama, and solar elevation angle alpha based on the vehicle solar vector S:
[0080] Based on the position of the solar vector in the inertial system in Step 1, describe the position of the solar vector in the vehicle orbital system using the solar altitude angle beta, solar azimuth angle gama, and solar elevation angle alpha, as the basis for subsequent attitude and sail design. The origin of the vehicle orbital system is the vehicle's center of mass, the OZ axis points to the center of the earth, the OX axis points perpendicular to the OZ in the orbital plane and in the vehicle's flight direction, and the OY axis is perpendicular to the OZ and OX axes according to the right-hand rule.
[0081] Make the following definitions:
[0082] Solar altitude angle beta: The angle between the solar vector in the vehicle orbital system and the YOZ plane of the vehicle orbital system.
[0083] Solar azimuth angle gama: The angle between the solar vector in the vehicle orbital system and the XOZ plane of the vehicle orbital system.
[0084] Solar elevation angle alpha: The angle between the solar vector in the vehicle orbital system and the XOY plane of the vehicle orbital system.
[0085] First, transform the inertial system solar vector S obtained in Step 1 i to the vehicle orbital system:
[0086] S o = A b←i * S i
[0087] Among them, A b←i is the attitude transformation matrix from the inertial coordinate system to the orbital coordinate system, and S o is the solar vector in the vehicle orbital system.
[0088] According to the definition:
[0089] alpha = arctan(-S z / S x )
[0090] gama = arctan(S y / S x )
[0091] beta = arcsin(S y / S o )
[0092] where, S x is the component of the solar vector in the vehicle orbit system on the X-axis of the vehicle orbit system, S y is the component of the solar vector in the vehicle orbit system on the Y-axis of the vehicle orbit system, S z is the component of the solar vector in the vehicle orbit system on the Z-axis of the vehicle orbit system.
[0093] Step 3: Set the yaw attitude maneuver angle and the sailboard rotation speed of the vehicle according to the solar angle
[0094] The design idea of the vehicle attitude maneuver angle, that is, the vehicle target attitude angle, is to complete the design of the optimal target attitude angle and the sailboard rotation speed under the following constraints:
[0095] Constraint condition 1: The vehicle always keeps its head facing the sun.
[0096] Constraint condition 2: The optimal incident angle of the solar sailboard is less than or equal to 20 degrees.
[0097] Constraint condition 3: The vehicle reduces attitude maneuvers as much as possible per orbit to ensure the vehicle's TT&C requirements.
[0098] Constraint condition 4: Under an inclined orbit, the range of the solar altitude angle beta changes within [-70, 70] in a year.
[0099] According to the above constraints and the three solar angles defined in Step 2, the method for setting the target attitude angle is as follows:
[0100] 1) When the solar altitude angle is [-90, -50], the target attitude angle is [0, 0, 90];
[0101] 2) When the solar altitude angle is (-50, -20] and the solar elevation angle is [-90, 90], the target attitude angle is [0, 0, 50];
[0102] 3) When the solar altitude angle is (-50, -20] and the solar elevation angle is [90, 180] or [-180, -90], the target attitude angle is [0, 0, 130];
[0103] 4) When the solar altitude angle is in the range of (-20, 20], the target attitude angle is [0, 0, 0];
[0104] 5) When the solar altitude angle is in the range of (20, 50] and the solar elevation angle is in the range of [90, 180] or [-180, -90], the target attitude angle is [0, 0, -130];
[0105] 6) When the solar altitude angle is in the range of (20, 50] and the solar elevation angle is in the range of [-90, 90], the target attitude angle is [0, 0, -50];
[0106] 7) When the solar altitude angle is in the range of [-90, -50), the target attitude angle is [0, 0, -90].
[0107] After the aircraft completes the attitude maneuver angle setting, the sailboard rotation speed is set. According to the aircraft attitude angle designed above, after the aircraft completes the attitude maneuver, it can ensure that the angle between the solar vector and the normal plane of the aircraft sailboard is minimized on average within one orbit. At this time, setting the sailboard rotation speed to make the sailboard normal follow the projection of the solar vector on the normal plane of the sailboard can ensure that the angle between the solar sailboard and the sun is minimized and the solar sailboard has the highest efficiency in facing the sun.
[0108] After the aircraft attitude maneuver is in place, the sailboard rotation speed is set. The sailboard rotation speed setting goes through two stages. The first stage is the capture stage, and the second stage is the tracking stage. The following definitions are made:
[0109] zeta is the deviation between the current angle and the desired angle of the sailboard. The desired angle of the sailboard is the sailboard angle that satisfies the condition that the sailboard normal follows the projection of the solar vector on the normal plane of the sailboard.
[0110] zeta = mod2pi(mod2pi(-alpha + pi))
[0111] w is the orbital angular velocity of the aircraft.
[0112] S ox is the component of the solar vector in the aircraft orbital system on the x-axis of the aircraft orbital system.
[0113] 1) If abs(beta) ≥ 50
[0114] The sailboard normal coincides with the X-axis of the local system.
[0115] 2) If 50 > abs(beta) > 20
[0116] If abs(zeta) > 5, then the sailboard capture is performed
[0117] If abs(zeta) ≤ 5, then the sailboard rotation speed is sign(S ox ) * w.
[0118] 3) If abs(beta) ≤ 20
[0119] If abs(zeta) > 5, then perform windsurfing board capture
[0120] If abs(zeta) ≤ 5, then the rotational speed of the windsurfing board is w.
[0121] The above-described embodiments are only relatively preferred specific embodiments of the present invention. Ordinary changes and substitutions made by those skilled in the art within the scope of the technical solution of the present invention should be included in the protection scope of the present invention.
Claims
1. A method for yaw maneuvering a tilted low-orbit aircraft towards the sun, characterized in that: The steps include: Step 1: Calculate the sun vector in the inertial system based on the time and spacecraft orbit information; Step 2: Calculate the solar altitude angle, solar azimuth angle and solar pitch angle in the spacecraft orbit system according to the solar vector; wherein the solar altitude angle is the angle between the solar vector in the spacecraft orbit system and the YOZ plane of the spacecraft orbit system, the solar azimuth angle is the angle between the solar vector in the spacecraft orbit system and the XOZ plane of the spacecraft orbit system, and the solar pitch angle is the angle between the solar vector in the spacecraft orbit system and the XOY plane of the spacecraft orbit system; Step 3: Based on the constraints and the solar altitude angle, solar azimuth angle and solar pitch angle in the spacecraft orbit system obtained in step 2, design the target attitude angle of the spacecraft yaw attitude maneuver; after the spacecraft maneuvers to the target attitude angle, design the windsurfing board rotation speed.
2. The method for yaw maneuvering of an inclined low-orbit aircraft towards the sun according to claim 1, characterized in that: The constraints of the aircraft yaw attitude maneuver are: Constraint 1: The aircraft always keeps its head facing the sun; Constraint 2: The optimal incident angle of the solar panel is less than or equal to 20 degrees; Constraint 3: The aircraft should minimize attitude maneuvers in each orbit to ensure the aircraft measurement and control requirements; Constraint 4: Under the inclined orbit, the solar altitude angle varies within the range of [-70,70] in one Earth year.
3. The method for yaw maneuvering of an inclined low-orbit aircraft towards the sun according to claim 2, characterized in that: Satisfying all constraints, the target attitude angle of the designed aircraft yaw attitude maneuver is: 1) When the sun altitude angle is [-90, -50], the target attitude angle is [0, 0, 90]; 2) When the solar altitude angle is (-50,-20] and the solar pitch angle is [-90,90], the target attitude angle is [0,0,50]; 3) When the solar altitude angle is (-50, -20] and the solar pitch angle is [90, 180] or [-180, -90], the target attitude angle is [0, 0, 130]; 4) When the sun altitude angle is (-20,20], the target attitude angle is [0,0,0]; 5) When the solar altitude angle is (20,50] and the solar pitch angle is [90,180] or [-180,-90], the target attitude angle is [0,0,-130]; 6) When the solar altitude angle is (20,50] and the solar pitch angle is [-90,90], the target attitude angle is [0,0,-50]; 7) When the sun altitude angle is [-90,-50), the target attitude angle is [0,0,-90].
4. The method for yaw maneuvering of an inclined low-orbit aircraft towards the sun according to claim 1, characterized in that: After the spacecraft maneuvers to the target attitude angle, the windsurfing rotation speed is designed to capture and track the windsurfing normal so that the windsurfing normal keeps up with the projection of the sun vector on the windsurfing normal surface, thus ensuring that the solar windsurfing angle to the sun is minimum and the solar windsurfing efficiency is highest.
5. The method for yaw maneuvering of an inclined low-orbit aircraft towards the sun according to claim 4, characterized in that: After the aircraft maneuvers to the target attitude angle, the windsurfing speed is designed as follows: 1) If abs(beta)≥50, the normal of the sailboard coincides with the X-axis of the system; where beta is the solar altitude angle; 2) If 50>abs(beta)>20, then: Calculate the deviation zeta between the current angle of the sailboard and the desired angle. The desired angle of the sailboard is the angle of the sailboard that satisfies the projection of the normal line of the sailboard to the normal surface of the sailboard. zeta=mod2pi(mod2pi(-alpha+pi)) Where alpha is the sun's pitch angle; If abs(zeta)>5, then sailboard capture is performed; If abs(zeta)≤5, the speed of the windsurfing board is sign(S ox )*w, for tracking; where w is the orbital angular velocity of the spacecraft, S ox is the component of the spacecraft orbit system sun vector on the x-axis of the spacecraft orbit system; 3) If abs(beta)≤20, then: Calculate the deviation zeta between the current angle of the windsurfing board and the desired angle; If abs(zeta)>5, then sailboard capture is performed; If abs(zeta)≤5, the speed of the windsurfing board is w and tracking is performed.
6. The method for yaw maneuvering of an inclined low-orbit aircraft towards the sun according to claim 1, characterized in that: According to the solar vector, the solar altitude angle, solar azimuth angle and solar pitch angle under the spacecraft orbit system are calculated, specifically: The sun vector S in the inertial system i Go to the spacecraft orbit system: S o =A b←i *S i In the formula, A b←i The attitude transformation matrix from the inertial coordinate system to the spacecraft orbit system, S o is the solar vector of the spacecraft orbit system; Then the solar altitude angle beta is: beta=arcsin(S y / S o ) The solar azimuth gama is: gama=arctan(S y / S x ) The solar pitch angle alpha is: alpha=arctan(-S z / S x ) In the formula, S x is the component of the solar vector of the spacecraft orbit system on the X-axis of the spacecraft orbit system, S y is the component of the solar vector of the spacecraft orbit system on the Y axis of the spacecraft orbit system, S z It is the component of the sun vector in the spacecraft orbit system on the Z axis of the spacecraft orbit system.
7. The method for yaw maneuvering of an inclined low-orbit aircraft towards the sun according to claim 1, characterized in that: According to the time and spacecraft orbit information, calculate the sun vector S in the inertial system i : Among them, Ω r u is the right ascension of the ascending node in the inertial system, which is 0; r is the latitude amplitude in the inertial system, i r is the orbital inclination in the inertial system.
8. The method for yaw maneuvering of an inclined low-orbit aircraft towards the sun according to claim 7, characterized in that: Use first-order quantities to calculate the solar vector in an inertial system: u r =ω r +f r oh r =(282.937347+0.32256206T JC )×(π / 180) <h2 style=";text-align:left;direction:ltr">f<h2 style=";text-align:left;direction:ltr"> r <h2 style=";text-align:left;direction:ltr"> =M<h2 style=";text-align:left;direction:ltr"> r <h2 style=";text-align:left;direction:ltr"> +2e<h2 style=";text-align:left;direction:ltr"> r <h2 style=";text-align:left;direction:ltr"> sin(M<h2 style=";text-align:left;direction:ltr"> r <h2 style=";text-align:left;direction:ltr"> )+1.25e<h2 style=";text-align:left;direction:ltr"> r <h2 style=";text-align:left;direction:ltr"> 2 <h2 style=";text-align:left;direction:ltr"> sin(2M<h2 style=";text-align:left;direction:ltr"> r <h2 style=";text-align:left;direction:ltr"> ) M r =(357.5291+35999.0502888889T JC )×(π / 180) I r =(23.439291-0.01300417T JC )×(π / 180) In the formula, ω r is the argument of perigee, a first-order quantity in rad; M r is the mean anomaly angle, modulo 2π, unit rad; T JC is the Julian century number from the current time to J2000.0; f r is the true anomaly angle, in rad.