JPS63173800A - Dual spin turn system - Google Patents
Dual spin turn systemInfo
- Publication number
- JPS63173800A JPS63173800A JP62003246A JP324687A JPS63173800A JP S63173800 A JPS63173800 A JP S63173800A JP 62003246 A JP62003246 A JP 62003246A JP 324687 A JP324687 A JP 324687A JP S63173800 A JPS63173800 A JP S63173800A
- Authority
- JP
- Japan
- Prior art keywords
- wheel
- dual spin
- satellite
- spacecraft
- axis
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
- 230000009977 dual effect Effects 0.000 title claims description 11
- 238000010586 diagram Methods 0.000 description 4
- 238000000034 method Methods 0.000 description 3
- 230000000694 effects Effects 0.000 description 2
- 230000001133 acceleration Effects 0.000 description 1
- 238000003780 insertion Methods 0.000 description 1
- 230000037431 insertion Effects 0.000 description 1
- 230000010354 integration Effects 0.000 description 1
Landscapes
- Preparation Of Compounds By Using Micro-Organisms (AREA)
Abstract
(57)【要約】本公報は電子出願前の出願データであるた
め要約のデータは記録されません。(57) [Summary] This bulletin contains application data before electronic filing, so abstract data is not recorded.
Description
【発明の詳細な説明】
〔産業上の利用分野〕
本発明は、宇宙飛翔体の姿勢をホイールの回転軸方向が
宇宙飛翔体の持つ角運動量方向に一致するように制御す
る自律制御型のデュアルスピンターン方式に関する。[Detailed Description of the Invention] [Industrial Application Field] The present invention is an autonomously controlled dual system that controls the attitude of a spacecraft so that the direction of the rotation axis of the wheel matches the direction of the angular momentum of the spacecraft. Regarding the spin-turn method.
以下弦日
〔従来の技術〕
従来のデュアルスピンターン方式は、宇宙飛翔体（衛星
）の姿勢を参照せずに、開ループ制御によりホイールを
加速することにより行われてきた。第２図に従来のデュ
アルスピンターン方式の原理図を示す。衛星本体７は第
２図い）に示すように、デュアルスピンターン前は角運
動量ベクトル８のまわ、９に回転し、ホイール９は静止
している。ここでホイール９を回転させることにより第
２図（ｂ） Ｋ示すように最終的に衛星本体７は慣性空
間に対して静止し、ホイール９のみが回転している状態
に移動する。[Prior Art] The conventional dual spin-turn method has been performed by accelerating the wheels by open-loop control without referring to the attitude of the spacecraft (satellite). FIG. 2 shows a diagram of the principle of the conventional dual spin-turn system. As shown in Figure 2), the satellite body 7 rotates around the angular momentum vector 8 and 9 before the dual spin turn, and the wheel 9 remains stationary. By rotating the wheel 9, the satellite main body 7 finally comes to a standstill with respect to the inertial space, with only the wheel 9 rotating, as shown in FIG. 2(b) K.
上述した従来のデュアルスピンターン方式では、ホイー
ルの回転軸が衛星の最大慣性主軸に一致している場合に
は、衛星本体は最終的に慣性空間に対して静止せずに、
デュアルスピンターンは失敗するという問題点がある。In the conventional dual spin-turn method described above, if the axis of rotation of the wheel coincides with the maximum principal axis of inertia of the satellite, the satellite body does not ultimately stand still with respect to inertial space.
The problem with dual spin turns is that they fail.
本発明は従来のもののこのような問題点を解決しようと
するもので、ホイールの回転軸か衛星の最大慣性主軸に
一致していてもデュアルスピンターンを遂行できるよう
にすることを目的とする。The present invention aims to solve the above-mentioned problems of the conventional technology, and aims to enable a dual spin turn to be performed even if the rotation axis of the wheel coincides with the maximum principal axis of inertia of the satellite.
本発明によれば、ホイールおよび慣性基準装置を有する
宇宙飛翔体において、前記慣性、・：基準装置の出力信
号を参照しながらホイールを駆動し、ホイールの回転軸
か宇宙飛翔体のもつ角運動量方向に一致するように衛星
の姿勢を制御することを特徴とするデュアルスピンター
ン方式％式％
次に９本発明の実施例について図面を参照して説明する
。第１図は本発明の一実施例の構成を示すブロック図で
ある。慣性基準装置１は衛星Ｙ軸のまわシの角速度信号
ω、を出力する。According to the present invention, in a spacecraft having a wheel and an inertial reference device, the wheel is driven while referring to the output signal of the inertia, reference device, and the rotational axis of the wheel is rotated in the direction of the angular momentum of the spacecraft. A dual spin-turn system characterized by controlling the attitude of the satellite so as to match the following.Next, nine embodiments of the present invention will be described with reference to the drawings. FIG. 1 is a block diagram showing the configuration of an embodiment of the present invention. The inertial reference device 1 outputs an angular velocity signal ω of rotation of the satellite Y-axis.
この信号は符号判定ロジック２および積分ロジック５を
通ってホイール４のサーボフィードバックループのトル
クコマンドＴとなる。トルクコマンドＴは減算器５によ
りホイール４の回転速度信号ωと減算処理が施され、ホ
イール４のサーボフィードバックループのゲイン６全通
してホイール４を駆動し、衛星のＹ軸のまわりにリアク
ションを発生する。This signal passes through the sign determination logic 2 and the integration logic 5 and becomes the torque command T of the servo feedback loop of the wheel 4. The torque command T is subtracted from the rotational speed signal ω of the wheel 4 by the subtractor 5, and the wheel 4 is driven through the entire gain 6 of the servo feedback loop of the wheel 4, generating a reaction around the Y axis of the satellite. do.
上記第１図の糸では、トルクコマンドＴは常にωＹπ≧
０（πはホイール回転加速度）となるように働くため１
次式によシ衛星は最終的に慣性空間に対して静止するこ
とが、Ｙ軸が最大慣性主軸であるかないかにかかわらず
保証される。For the thread shown in Figure 1 above, the torque command T is always ωYπ≧
0 (π is the wheel rotational acceleration), so it is 1
According to the following formula, it is guaranteed that the satellite will finally come to rest with respect to inertial space, regardless of whether or not the Y-axis is the maximum principal axis of inertia.
■（衛星本体のもつ運動エイ・ルギー）、 〔発明の効
果〕
以上説明したように０本発明はホイールの回転軸が衛星
の最大慣性主軸に一致していても。(Motion of the satellite main body) [Effects of the invention] As explained above, the present invention can be used even if the axis of rotation of the wheel coincides with the principal axis of maximum inertia of the satellite.
デュアルスピンターンを遂行できるので、宇宙飛翔体の
姿勢制御に利用できる効果がある。特に、ロケットのノ
ーズフェアリングの幾何形状の制約やミッション機器の
形状の制御等から初期軌道投入方向を最大慣性主軸にす
ることのできない特殊ミッション用衛星に利用できる点
。Since it can perform dual spin turns, it can be used to control the attitude of spacecraft. In particular, it can be used for special mission satellites where the initial orbit insertion direction cannot be set as the maximum principal axis of inertia due to constraints on the geometry of the rocket's nose fairing or control of the shape of mission equipment.
その得られる効果は太きい。The effects obtained are profound.
第１図は本発明の一実施例の構成を示すブロック図、第
２図（ａ） （１））は従来例の動作を説明する図であ
る。
図において、１は慣性基準装置、２は符号判のゲイン、
７は衛星本体である。
第１図
第２図
（ａ） （ｂ）
ｌ衛鳳本体FIG. 1 is a block diagram showing the configuration of an embodiment of the present invention, and FIG. 2(a) (1)) is a diagram explaining the operation of a conventional example. In the figure, 1 is an inertial reference device, 2 is a code size gain,
7 is the satellite body. Figure 1 Figure 2 (a) (b) Main body
Claims (1)
て、前記慣性基準装置の出力信号を参照しながら前記ホ
イールを駆動し、該ホイールの回転軸が宇宙飛翔体自体
のもつ角運動量方向に一致するように該宇宙飛翔体の姿
勢を制御することを特徴とするデュアルスピンターン方
式。1. In a spacecraft having a wheel and an inertial reference device, drive the wheel while referring to the output signal of the inertial reference device so that the axis of rotation of the wheel coincides with the direction of the angular momentum of the spacecraft itself. A dual spin-turn system characterized by controlling the attitude of the spacecraft.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP62003246A JPS63173800A (en) | 1987-01-12 | 1987-01-12 | Dual spin turn system |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP62003246A JPS63173800A (en) | 1987-01-12 | 1987-01-12 | Dual spin turn system |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPS63173800A true JPS63173800A (en) | 1988-07-18 |
Family
ID=11552100
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP62003246A Pending JPS63173800A (en) | 1987-01-12 | 1987-01-12 | Dual spin turn system |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPS63173800A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH01117097A (en) * | 1987-10-29 | 1989-05-09 | Sanyo Electric Co Ltd | Feed mechanism for electronic component |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS5686900A (en) * | 1979-12-17 | 1981-07-15 | Mitsubishi Electric Corp | Controlling system for number of revolution of artificial satellite using rotary wheel |
| JPS59143799A (en) * | 1983-02-07 | 1984-08-17 | 三菱電機株式会社 | Controller for attitude of artificial satellite |
-
1987
- 1987-01-12 JP JP62003246A patent/JPS63173800A/en active Pending
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS5686900A (en) * | 1979-12-17 | 1981-07-15 | Mitsubishi Electric Corp | Controlling system for number of revolution of artificial satellite using rotary wheel |
| JPS59143799A (en) * | 1983-02-07 | 1984-08-17 | 三菱電機株式会社 | Controller for attitude of artificial satellite |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH01117097A (en) * | 1987-10-29 | 1989-05-09 | Sanyo Electric Co Ltd | Feed mechanism for electronic component |
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