CN102874418B - Method for improving orbit-transferring safety of inclined orbit satellite - Google Patents

Abstract

A method to improve the safety of orbit change of inclined orbit satellites is to consider the sun altitude angle condition during the design process of the flight program, and through reasonable adjustment of flight events, it solves the problem of the orbit change of inclined orbit satellites under the condition of high sun angle. The problem of energy shortage during maneuvering can also simplify the flight control operation. Designing the flight program according to the method can shorten the battery discharge time during the orbit change ignition period, reduce the load power during the battery power supply period, thereby reducing the battery discharge depth, ensuring the energy security of the satellite, and making the launch window not limited by the sunlight angle.

Description

A Method of Improving the Safety of Orbit Changing for Inclined Orbit Satellites

technical field

The invention relates to a method for improving the safety of orbit change of an inclined orbit satellite, which can ensure the safety of energy and thermal control during satellite orbit change, and is mainly used during flight control of inclined orbit satellites.

Background technique

Generally, the geosynchronous orbit satellite has a low sun altitude angle after launch into orbit, and the sun illumination angle (the angle between the sun's rays and the projection of the sun's rays on the satellite's orbital plane) is small during orbit change ignition, which meets the power supply demand of the satellite. Therefore, the flight program design during the orbit change generally does not need to consider the problem of energy shortage. The common flight program design method during the orbit change is: according to the orbit change strategy and the tracking arc section, arrange the earth capture sequentially and establish the earth pointing attitude, the earth pointing Measurement and control events such as gyro calibration under attitude, preparation before orbit change, sail board and establishment of ignition attitude, apogee ignition, status setting after ignition, etc.

However, for satellites in inclined orbits, when launched in different seasons, the solar elevation angle varies greatly. In the case of high sun angle, during the whole period from the establishment of the ignition attitude of the satellite to the restoration of the earth pointing attitude after the ignition, the sun illumination angle (the angle between the normal line of the sail and the sunlight) is relatively large, and the output power of the solar wing cannot Satisfy the demand for satellite power supply. When the sun’s altitude angle reaches above 70°, the output power is very small, and it is mainly powered by the battery. The battery discharge time is from the time when the satellite establishes an orbital ignition attitude to the time when the sailboard is re-aligned to the sun after the ignition. The load power includes the satellite The power of platform equipment and associated heaters. According to calculations, in the case of high sun angles, if the common flight program design method for medium and high orbit satellites is adopted, the discharge depth of the battery will exceed the design value, which will affect the power supply safety of the satellite. If the design of the satellite is changed, such as increasing the battery capacity, or changing the launch window of the satellite, such as only choosing to launch in the season with a small solar altitude angle, it may increase the development cost of the satellite, and even affect the completion of the satellite launch task. Network satellites may not be able to avoid launching during high sun angle seasons.

Contents of the invention

The technical problem of the present invention is: to overcome the deficiencies of the prior art, and to provide a method for improving the safety of orbit change of an inclined orbit satellite under the condition of high sun angle, by reasonably adjusting the temperature control threshold of the heater, the satellite attitude and the vertical sail The implementation timing of flight events such as the board, as far as possible to reduce the discharge depth of the battery during the orbit change, to ensure the safety of energy and thermal control of the satellite.

The technical solution of the present invention is: a kind of method for improving the safety of orbit change of inclined orbit satellite, the steps are as follows:

(1) Calculate the solar altitude angle θs at the ignition moment t, and the solar altitude angle is the angle between the line connecting the earth and the sun and the orbital plane; if θs<y, then change orbit according to the conventional flight control procedure; otherwise Go to step (2), said y=arccos (load power/solar wing output full power at t moment)

(2) According to the time of sending the remote control command to establish the ignition attitude and the adjustment time of the yaw attitude, design the moment of establishing the ignition attitude;

(3) Calculating the solar illumination angle θsa0 of the earth pointing attitude moment t0 and establishing the solar illumination angle θsa1 of the ignition attitude moment t1 under the situation of not erecting the solar sail; the included angle;

(4) According to the calculation result of step (3), if θsa0<θsa1, then the sailboard operation is performed immediately after the satellite establishes the earth pointing attitude; otherwise, the sailboard operation is performed before the ignition attitude is established;

(5) Calculate the discharge depth X of the battery during the orbit change ignition period. If X>m, then after the earth pointing attitude is established, the temperature control threshold of the surface heater is not controlled to the upper limit of the safe value during the orbit change period, and the ignition attitude is established. Lower the temperature control threshold of the non-irradiated heater to the lower limit of the safety value; otherwise, do not adjust the temperature control threshold of the heater, and the m is the maximum allowable discharge depth of the transfer track battery;

(6) After judging whether the orbital perigee height is less than the minimum operating height requirement of the earth sensor on the satellite after the ignition is completed, if it is less than, the satellite turns into a sun-oriented cruising attitude before leaving the measurement, control and tracking arc; otherwise, turn to step ( 7);

(7) Design the satellite yaw angle offset and sailboard rotation angle under the earth pointing attitude: if the sun altitude angle θs calculated in step (1)<(90°-β), then set the yaw angle offset to - 90°, the sailboard rotation angle is -90°; otherwise, the yaw angle offset remains 0°, and the sailboard tracks the sun; the β=arccos (load power/full output power of the solar wing).

Battery discharge depth X=100%*((((P1-P0)*T1)/U)/Ah) in the described step (5); Wherein, P1 is the load power during orbit change ignition; Solar wing output power P0=P*cos(θsa), P is the full output power of the solar wing, θsa is the sunlight angle at the moment of ignition; T1 is the time from establishing the ignition attitude to the end of orbit change ignition and returning to the earth pointing attitude; U is the battery power supply The bus voltage; Ah is the rated ampere-hour of the battery.

Heat the gas cylinder to a temperature above Tw before orbit change ignition, Tw=tw+v*T2;

Among them, tw is the lower limit of the required value of the gas cylinder temperature; v is the temperature drop rate of the gas cylinder during the ignition process; T2 is the ignition time of orbit change.

After described step (6) ignites, increase and judge whether the sun elevation angle of satellite's next orbit change ignition moment t exceeds the measuring range of the sun sensor, if exceed, then judge that the earth pointing attitude is established before next orbit change ignition Whether the time until the ignition attitude is established satisfies the calibration time of the satellite gyroscope, if not, add a gyroscope calibration operation after the end of this orbit change.

The beneficial effect of the present invention compared with prior art is:

(1) The energy security of the satellite is the basis for the successful launch and normal operation of the satellite. However, when satellites in inclined orbits are launched in certain seasons, the altitude angle of the sun is high, which leads to the need for battery power during the entire time period from the establishment of the ignition attitude to the restoration of the earth pointing attitude after the ignition. If the battery discharge depth exceeds the design value, it will Affect satellite energy security. In the present invention, without changing the design of the satellite (such as increasing the battery capacity) or adjusting the launch window, through the reasonable adjustment of the temperature control threshold of the thermal control system of the satellite before and after the orbit change, the implementation time and sequence of flight events, and the attitude and sailboard rotation angle, It can ensure the energy and thermal control safety of the satellite under the condition of high sun altitude angle. On the one hand, it reduces the development cost of changing the satellite design, and at the same time reduces the restrictions on the launch window, making the launch time of inclined orbit satellites more flexible.

(2) The time interval between orbit changes of inclined orbit satellites is long, and there are arcs that are invisible to measurement and control. After each orbit change, the safety mode of sun orientation can ensure the safety of energy and thermal control of satellites. However, if it is switched to the sun-oriented safety mode, before the next orbit change, it is necessary to re-perform the measurement and control operations of the earth capture and the establishment of the earth pointing attitude. The invention provides a method for adjusting the yaw attitude and sailboard rotation angle of the satellite under the earth pointing attitude, which can simplify the measurement and control operation on the basis of ensuring the safety of satellite energy and thermal control.

(3) During the ignition process of orbit change, the temperature of the gas cylinder will drop rapidly. If the cylinder temperature is too low, it will affect the normal operation of the propulsion system. However, if the temperature of the gas cylinder is too high, the gas cylinder may explode. The invention can ensure that the temperature of the gas cylinder is within the range of the required value during the orbit change by rationally adjusting the temperature of the gas cylinder.

(4) After the satellite establishes the ignition attitude, if the field of view of the sun sensor is not satisfied, the gyro can only be used as the measurement reference for the yaw attitude, and the time for gyro calibration is required to be more than 40 minutes. Limits, resulting in insufficient gyro calibration time and inaccurate calibration results. The present invention can solve this problem by making reasonable use of the time after the track change ends.

Description of drawings

Figure 1 is a schematic diagram of the satellite orbit change process;

Fig. 2 is a flowchart of the present invention.

Detailed ways

Specific embodiments of the present invention will be further described in detail below in conjunction with the accompanying drawings.

The safety of energy and thermal control is the basis for the implementation of satellite orbit change maneuvers. However, for satellites in inclined orbits launched in some seasons (the common inclination angle is 55°), the solar altitude angle will be very large (if the inclination angle is 55°, the maximum solar altitude angle is 55° (inclination angle) + 23.5° (ecliptic and equatorial) ) = 78.5°), so that during the period from the establishment of the ignition attitude of the satellite to the restoration of the earth pointing attitude after the ignition, the sun illumination angle (the angle between the normal of the sail and the sunlight) is relatively large, and the solar wing The output power cannot meet the demand for satellite power supply, and battery power is required. However, the capacity of the battery is limited. If the depth of discharge is too large, it will affect the safety of the satellite's power supply.

The invention provides a method, which can greatly shorten the discharge time of the storage battery and reduce the load power by reasonably adjusting the temperature control threshold of the heater and the time and sequence of flight events, thereby reducing the discharge depth of the storage battery and ensuring the safety of satellite energy.

In order to understand the present invention more clearly, below at first the conventional orbit changing process of satellite is described, as shown in Figure 1, satellite generally carries out orbit changing 3-4 times, and each orbit changing process is similar, and main difference is that if changing orbit Before the ignition, the satellite is already working at the earth pointing attitude, so there is no need to carry out the earth capture operation.

Before the satellite changes orbit and ignites, it first captures the earth and establishes the earth pointing attitude (the OY axis is the pitch axis perpendicular to the orbital plane of the flight, and the positive direction points to the south pole of the celestial sphere; the OZ axis is the yaw axis positively points to the center of the earth; the OX axis is The rolling axis forms a right-handed spiral direction with the OY axis and the OZ axis). This is not done if the satellite is already in the Earth pointing attitude. Carry out pre-ignition preparations such as gyro calibration in the ground-pointing attitude, erect the sailboard (make the normal of the sailboard parallel to the X-axis of the satellite), establish the ignition attitude (the OX-axis is located in the orbital plane of the satellite flight, and the positive direction points to the satellite’s The forward direction; the OZ axis is positively pointing to the center of the earth; the OY axis, the OZ axis and the OX axis form a right-handed spiral direction). After the satellite orbit change ignition is completed, the earth pointing attitude is restored.

On the basis of the conventional orbit change process, the present invention improves the safety of the IGSO satellite orbit change through the following steps, which is mainly reflected in three aspects:

1. Optimization of flight events before orbit change ignition

(1) Calculate the sun altitude angle θs at the ignition moment t, the sun altitude angle is the angle between the sun ray vector and the orbital plane, if θs<y, then perform orbit change according to the conventional flight control program; otherwise go to step (2 ); said y=arccos (load power/solar wing output full power at t moment);

According to the orbit change strategy file of the satellite, the orbit of the satellite before and after each orbit change ignition can be known. Assuming that the components of the solar unit vector at the ignition time t in the orbital coordinate system are Sox, Soy, and Soz according to the orbital information, the solar altitude angle at the time t is: θs=arcsin(Soy). The above-mentioned θs<y degree is an empirical formula. In the case of θs<y degree, after the satellite establishes the ignition attitude, the output power of the solar wing can meet the load requirements, the battery is not discharged or the discharge is very small, and the flight control program does not need to consider the lack of energy. The problem.

(2) According to the time for sending the remote control command for establishing the ignition attitude and the time for adjusting the yaw attitude, design and establish the ignition attitude time t2;

Start the operation of establishing the ignition attitude t2 minutes before switching to the ignition mode. In the case of high solar altitude, in order to reduce the battery discharge time, it is necessary to delay the time t2 as much as possible, which can be set as follows: t2=(send to establish ignition Attitude remote command time + yaw attitude adjustment time (including attitude stabilization time))*2.

(3) Calculating the solar illumination angle θsa0 of the earth pointing attitude moment t0 and establishing the solar illumination angle θsa1 of the ignition attitude moment t1 under the situation of not erecting the solar sail; the included angle;

Assume that the components of the sun unit vector in the satellite body coordinate system at t0 and t1 calculated according to the orbit and attitude information are Sbx, Sby, and Sbz, and the solar illumination angle is calculated according to the specific body orientation of the solar panel. If the solar panel points to the -X axis , then the solar illumination angle is: θsa0=arccos(-Sbx); if the solar panel points to the -Z axis, the solar illumination angle is: θsa1=arccos(-Sbz).

(4) According to the calculation result of step (3), if θsa0<θsa1, then the sailboard operation is performed immediately after the satellite establishes the earth pointing attitude; otherwise, the sailboard operation is performed before the ignition attitude is established;

(5) Calculate the discharge depth X of the battery during the orbit change ignition period. If X>m, then after the earth pointing attitude is established, the temperature control threshold of the surface heater is not controlled to the upper limit of the safe value during the orbit change period, and the ignition attitude is established. Lower the temperature control threshold of the non-irradiated heater to the lower limit of the safety value; otherwise, do not adjust the temperature control threshold of the heater, and the m is the maximum allowable discharge depth of the transfer track battery;

Battery discharge depth X=100%*((((P1-P0)*T1)/U)/Ah); among them, P1 is the load power during orbit change ignition; solar wing output power P0=P*cos(θsa) , P is the full output power of the solar wing, θsa is the solar illumination angle at the ignition moment; T1 is the time from establishing the ignition attitude to the end of orbit change ignition and returning to the earth pointing attitude; U is the bus voltage under the condition of battery power supply; Ah is the battery voltage Rated ampere hours. ;

2. Satellite attitude setting after orbit change

Before the satellite reaches the target orbit, the satellite can turn to the sun-oriented cruise attitude or the ground-pointing attitude after each orbit change. By adjusting the yaw attitude and the sail angle, the sun illumination angle can always be kept at the solar wing output. Range to meet load requirements. Setting the attitude of the satellite according to the following method can meet the energy demand and simplify the measurement and control operation at the same time:

(6) If the perigee height of the orbit after the orbit change ignition is less than the minimum altitude requirement of the earth sensor, the satellite will turn to the sun-oriented cruise attitude before leaving the measurement, control and tracking arc; otherwise, go to step (7) ;

(7) Design the satellite yaw angle offset and sailboard rotation angle under the earth pointing attitude: if the sun altitude angle θs calculated in step (1)<(90°-β), then set the yaw angle offset to - 90°, so that the satellite +Y points to the forward direction of the satellite (assuming that the X plane is illuminated by the long-term on-orbit design condition of the satellite), the sail panel rotation angle is -90°, which can ensure that the output power of the solar wing meets the load requirements; otherwise, The yaw angle offset remains unchanged at 0°, and the sailboard tracks the sun (when the sun illumination angle>β, the sailboard rotation angle needs to be adjusted, which is a well-known practice in the industry and will not be described in detail); the β=arccos(load power /Sunwing output full power).

3. Other security measures

1. During orbital ignition, the temperature of the gas cylinder will drop rapidly. In order to ensure that the temperature of the gas cylinder is within the required value range, it is necessary to heat the gas cylinder to a certain temperature above Tw before ignition, Tw=tw+v*T2;

Among them, tw is the lower limit of the required value of the gas cylinder temperature; v is the temperature drop rate of the gas cylinder during the ignition process; T2 is the ignition time of orbit change.

2. When the sun altitude angle exceeds the measurement range of the sun sensor (generally 60°), it is necessary to use a gyro to measure the yaw attitude of the satellite. In order to ensure the accuracy of the satellite's orbit change, it is necessary to accurately calibrate the gyro. The timing of gyro calibration is generally arranged within the time period T3 before the ignition attitude is established and stabilized to the establishment of the ignition attitude. Since the gyro calibration time is generally required to be more than 40 minutes, if T3 < 40 minutes before the next orbit change ignition, then in this After the first orbit change, add a gyro calibration operation, which will be used for the next orbit change.

In the actual application process, the method of the present invention can adjust the corresponding content of the flight control program according to the contents of the above-mentioned parts on the basis of the conventional flight control program, so as to achieve the purpose of improving the safety of satellite orbit change. During the orbit change of many inclined orbit satellites in our country, the invention greatly reduces the discharge depth of the battery during the orbit change, simplifies the flight control operation, and ensures that the energy and heat of the satellite orbit change maneuvering process is ensured by the temperature of the gas cylinder within the required value range. Control security.

The content that is not described in detail in the specification of the present invention belongs to the well-known technology of those skilled in the art.

Claims (3)

1. A method for improving the safety of orbit changing of inclined orbit satellites is characterized in that the steps are as follows: (1) Calculate the solar altitude angle θs at the ignition moment t, and the solar altitude angle is the angle between the line connecting the earth and the sun and the orbital plane; if θs<y, then change orbit according to the conventional flight control procedure; otherwise Go to step (2), said y=arccos (load power/solar wing output full power at t moment); (2) According to the time of sending the remote control command to establish the ignition attitude and the adjustment time of the yaw attitude, design the moment of establishing the ignition attitude; (3) Calculating the solar illumination angle θsa0 of the earth pointing attitude moment t0 and establishing the solar illumination angle θsa1 of the ignition attitude moment t1 under the situation of not erecting the solar sail; the included angle; (4) According to the calculation result of step (3), if θsa0<θsa1, then the sailboard operation is performed immediately after the satellite establishes the earth pointing attitude; otherwise, the sailboard operation is performed before the ignition attitude is established; (5) Calculate the discharge depth X of the battery during the orbit change ignition period. If X>m, then after the earth pointing attitude is established, the temperature control threshold of the surface heater is not controlled to the upper limit of the safe value during the orbit change period, and the ignition attitude is established. Before lowering the temperature control threshold of the non-irradiated heater to the lower limit of the safety value; otherwise, the temperature control threshold of the heater is not adjusted, and the m is the maximum allowable discharge depth of the transfer track battery; the battery discharge depth X = 100%*((((P1-P0)*T1)/U)/Ah); among them, P1 is the load power during orbit change ignition; solar wing output power P0=P*cos(θsa), P is the solar wing Output full power, θsa is the sun illumination angle at the ignition moment; T1 is the time from establishing the ignition attitude to the end of orbit change ignition and returning to the earth pointing attitude; U is the bus voltage under the condition of battery power supply; Ah is the rated ampere-hour of the battery; (6) After judging whether the orbital perigee height is less than the minimum operating height requirement of the earth sensor on the satellite after the ignition is completed, if it is less than, the satellite turns into a sun-oriented cruising attitude before leaving the measurement, control and tracking arc; otherwise, turn to step ( 7); (7) Design the satellite yaw angle offset and sailboard rotation angle under the earth pointing attitude: if the sun altitude angle θs calculated in step (1)<(90°-β), then set the yaw angle offset to - 90°, the sailboard rotation angle is -90°; otherwise, the yaw angle offset remains 0°, and the sailboard tracks the sun; the β=arccos (load power/full output power of the solar wing). 2. A kind of method for improving the safety of orbit change of inclined orbit satellite according to claim 1, is characterized in that: before the orbit change ignition, the gas cylinder is heated to more than the temperature Tw, Tw=tw+v*T2; Among them, tw is the lower limit of the required value of the gas cylinder temperature; v is the temperature drop rate of the gas cylinder during the ignition process; T2 is the ignition time of orbit change. 3. according to claim 1 or 2, a kind of method for improving the safety of orbit change of satellite in inclined orbit is characterized in that: after described step (6) ignites, increase the sun for judging the next time t of orbit change ignition of satellite Whether the altitude angle exceeds the measurement range of the sun sensor, if it exceeds, then judge whether the time from establishing the earth pointing attitude before the next orbit change ignition to establishing the ignition attitude meets the calibration time of the satellite gyro, if not, then in this Add a gyro calibration operation after the first orbit change.