地球同步卫星相关的专利数据

序号	专利名	申请号	申请日	公开（公告）号	公开（公告）日	发明人
61	Sequenced heat rejection for body stabilized geosynchronous satellites	US203710	1998-12-02	US06073888A	2000-06-13	Walter S. Gelon; John C. Hall; Christian J. Goodman
In one embodiment, a thermal radiative system for an earth orbiting satellite including a plurality of faces intermittently exposed to maximum solar illumination comprises a thermal radiator is mounted on a face for discharging heat from a thermal load to deep space. A heat conductor extends between the thermal load and the thermal radiator. Thermal switches are operable for connecting the thermal load to the thermal radiator for cooling when the temperature of the thermal load is above a predetermined level and for disconnecting the thermal load from the thermal radiator when the temperature of the thermal load falls below the predetermined level. A shield including phase change management (PCM) material is thermally connected to the thermal load for drawing heat away therefrom. In another embodiment, first and second thermal radiators are mounted on first and second faces, respectively, of the satellite for discharging heat from the thermal load and the thermal load is sequentially connected to the first and second thermal radiators to achieve optimum discharge of heat from the thermal load. To this end, the thermal load is thermally connected to the first radiator so long as the first radiator is exothermic and thermally disconnected from the first radiator when the first radiator becomes endothermic and is thermally connected to the second radiator so long as the second radiator is exothermic and thermally disconnected from the second radiator when the second radiator becomes endothermic.
62	Sequenced heat rejection for body stabilized geosynchronous satellites	EP99309581.9	1999-11-30	EP1006769A3	2000-11-22	Gelon, Walter S.; Goodman, Christian J.; Hall, John C.
In one embodiment, a thermal radiative system (36) for an earth orbiting satellite including a plurality of faces (24-34) intermittently exposed to maximum solar illumination comprises a thermal radiator (38) is mounted on a face for discharging heat from a thermal load (40) to deep space. A heat conductor extends between the thermal load and the thermal radiator. Thermal switches (44) are operable for connecting the thermal load to the thermal radiator for cooling when the temperature of the thermal load (40) is above a predetermined level and for disconnecting the thermal load from the thermal radiator when the temperature of the thermal load falls below the predetermined level. A shield including phase change management (PCM) material is thermally connected to the thermal load for drawing heat away therefrom. In another embodiment, first and second thermal radiators are mounted on first and second faces, respectively, of the satellite for discharging heat from the thermal load and the thermal load is sequentially connected to the first and second thermal radiators to achieve optimum discharge of heat from the thermal load. To this end, the thermal load is thermally connected to the first radiator so long as the first radiator is exothermic and thermally disconnected from the first radiator when the first radiator becomes endothermic and is thermally connected to the second radiator so long as the second radiator is exothermic and thermally disconnected from the second radiator when the second radiator becomes endothermic.
63	Practical orbit raising system and method for geosynchronous satellites	US09328911	1999-06-09	US07113851B1	2006-09-26	Walter Gelon; Ahmed Kamel; Darren Stratemeier; Sun Hur-Diaz
A practical orbit raising method and system wherein a satellite quickly escapes the Van Allen radiation belts and payload mass and mission life are maximized. A satellite is launched that contains high thrust chemical propulsion thrusters, high specific impulse electric propulsion thrusters and a solar array. The satellite quickly escapes the Van Allen radiation belts by firing the high thrust chemical propulsion thrusters at apogees of intermediate orbits, starting from the transfer orbit initiated by a launch vehicle, to successively raise the perigees until the perigee clears the Van Allen radiation belts. The payload mass and mission life are maximized by firing high specific impulse electric propulsion thrusters to raise the satellite to near synchronous orbit, while steering the thrust vector and solar array to maintain the sun's illumination on the solar array. The chemical and/or electric propulsion thrusters are then fired to achieve geosynchronous orbit.
64	静止衛星用アンテナの方位角、仰角、偏波角、及び静止衛星方向表示携帯端末装置	JP2009004140	2009-05-28	JP3155170U	2009-11-12	鈴木　良幸; 良幸鈴木
【課題】静止衛星用アンテナの方向調整作業を行う場合、予めアンテナの方位角などの情報を準備することなく、任意の場所における目標静止衛星の方位角などの各値を表示すると共に、目標静止衛星の方向を矢印などで示すことによって、静止衛星用アンテナの方向調整作業を容易にすることが可能となる携帯端末装置を提供する。【解決手段】GPS衛星などの測位衛星受信部2と、方位検知部5と、静止衛星の経度情報データベースなどが保存される記憶部3と、情報演算部4と、表示部6などから構成される携帯端末装置であり、測位衛星で得られた現在地上位置の緯度・経度情報と、目標とする静止衛星の経度情報とから、現在地上位置における目標静止衛星のアンテナの方位角、仰角、偏波角の各値を表示し、磁北又は真北を基準とした静止衛星の方向を矢印などで表示することができる。【選択図】図2
65	Elliptical satellite system which emulates the characteristics of geosynchronous satellites	US09892132	2001-06-25	US06678519B2	2004-01-13	David Castiel; John Draim; Kenneth F. Manning
An elliptical satellite communication system including a constellation of satellites which orbit the earth at a height less than that necessary for geosynchronous orbits but which simulate the characteristics of geosynchronous orbits. The satellites' velocity near the apogee portion of their orbit approximates the rotational velocity of the earth, and during that period appear to hover over the earth. The ground stations on the earth always communicate with a satellite at or near its apogee, and hence that satellite appears to the ground station to hover over the earth. During the times when the satellite is outside the apogee portion, its communication is shut off to prevent any possibility of interfering with geosynchronous satellites and its power supply is used to charge a battery on the satellite. Thus, the power supply of the system can be reduced by an amount equivalent to the percentage of time the satellite is not used.
66	Elliptical satellite system which emulates the characteristics of geosynchronous satellites	US09056051	1998-04-06	US06263188B1	2001-07-17	David Castiel; John Draim; Kenneth F. Manning
An elliptical satellite system which carries out communication. The satellite orbits a height above the earth less than that necessary for geosynchronous orbits. When the satellite is near the apogee portion of its orbit, its velocity approximates the rotational velocity of the earth, and during that period it appears to hover over the earth. Each ground station on the earth always communicates a satellite within a predetermined position of its apogee, and hence that satellite appears to the ground station to hover over the earth. The satellite hence does not communicate with any earth station when it is outside of that apogee portion. During the times when the satellite is outside the apogee portion, its communication is therefore shut off to prevent any possibility of interfering with geosynchronous satellites. During this time, the power supply on the satellite is also used to charge a battery on the satellite. This enables the power supply to be made smaller by an amount equivalent to the duty cycle of the satellite: during the time which it is on.
67	System and method for the acquisition of a non-geosynchronous satellite signal	US949989	1997-10-14	US5929808A	1999-07-27	Amer A. Hassan; Sami M. Hinedi; James R. Miller
An antenna having multiple antenna elements that are selectively activated performs an initial acquisition of a satellite signal by activating only a few of the antenna elements. This results in a broad beam width and increases the likelihood of detecting the signal from the satellite. When the satellite signal is initially detected, the system increases the number of active elements to narrow the beam width. The system incrementally increases the number of active antenna elements until the detected signal from the satellite exceeds a predetermined threshold. At that point, the location of the satellite may be precisely determined and all antenna elements activated to lock onto the satellite signal. If the antenna loses acquisition of the satellite signal, the reverse process may be implemented whereby some elements are selectively deactivated to broaden the beam width of the antenna in an effort to reacquire the satellite signal. When the satellite signal is reacquired, the antenna elements are incrementally reactivated until all antenna elements are active.
68	System for controlling the direction of the momentum vector of a geosynchronous satellite	US118847	1980-02-05	US4325124A	1982-04-13	Udo Renner
A system for compensating the disturbance torques applied to a satellite, which eliminates the requirement for a thruster control loop. The disturbance torque itself is used as the compensating torque in order to super-impose to the incidental misalignment of the solar panel arrays an artificial misalignment that can cause the momentum vector to be adjusted to the desired direction in order to restore the correct attitude of the satellite. The direction of the momentum vector is controlled in orbit only by solar sailing, that is by organizing at prescribed times suitable manoeuvres of one of the solar panels in order to adjust the solar panel array configuration when the roll angle of the satellite exceeds a determined threshold value.
69	Free return lunar flyby transfer method for geosynchronous satellites	US57938	1998-04-09	US06116545A	2000-09-12	Jeremiah O. Salvatore; Cesar A. Ocampo
A method is provided for using a lunar flyby maneuver to transfer a satellite from a quasi-geosynchronous transfer orbit having a high inclination to a final geosynchronous orbit having a low inclination. The invention may be used to take the inclination of a final geosynchronous orbit of a satellite to zero, resulting in a geostationary orbit, provided that the satellite is launched in March or September.
70	정지위성의 임무분석을 통한 임무일정표 자동 생성 방법	KR1019960057783	1996-11-26	KR1019980038854A	1998-08-17	김재훈; 이정숙; 김재명
1. 청구 범위에 기재된 발명이 속한 기술분야 정지위성의 임무분석을 통한 임무일정표 자동 생성 방법 2. 발명이 해결하려고 하는 기술적 과제 각종 이벤트를 자동으로 분석하고 임무일정을 자동으로 생성하며, 임무에 관련된 운용일정과 원격명령 화일명을 생성하여 임무일정표를 자동으로 생성하는 임무일정표 자동 생성 방법을 제공하고자 함. 3. 발명의 해결방법의 요지 식과 센서 간섭현상 예측 데이타, 태양 간섭현상 예측 데이타, 및 궤도조정 데이타를 구하는 제 1 단계, 사용자 지정 이벤트에 따라 궤도조정 계획 이벤트, 위성추적 계획 이벤트, 및 궤도 결정 이벤트를 처리하는 제 2 단계, 자동처리 이벤트를 처리하는 제 3 단계, 및 이벤트 처리 결과 데이타와 경험 지식베이스를 이용하여 임무를 분석한 후에 임무일정, 운용일정, 및 위성으로 전송할 원격명령 화일명을 생성하여 임무일정표를 생성하는 제 4 단계를 포함한다. 4. 발명의 중요한 용도 위성관제시스템에 이용됨.
71	PROCEDE DE RECTIFICATION EN TEMPS REEL D'IMAGES DE SATELLITES METEOROLOGIQUES GEOSTATIONNAIRES	EP90916761.1	1990-10-31	EP0452461B1	1996-03-06	DE WAARD, Johannes; ADAMSON, Jan; BOS, Albert, Martinus
A method of real time correction of image data remotely detected by a radiation measuring instrument such as is used on rotationally stable geostationary satellites. The method comprises (1) predicting the correction parameters by a statistical analysis based on a number of previous images, (2) reading a limited number of raw image lines in order to facilitate the accurate setting of the South horizon or the North horizon, (3) determining the satellite's rotating speed from a limited number of raw image lines, (4) redefining the measuring instrument's orientation and aim parameters by using the horizon information obtained in real time, (5) computing the position of the other horizon and of the centre of the image on the basis of the horizon already detected, (6) creating a deformation matrix immediatly after determining the first horizon, and (7) correcting in real time the data of the image received immediately afterwards.
72	Antenna system for receiving signals that are transmitted by geostationary satellite	US09584102	2000-05-31	US06504504B1	2003-01-07	Daniel G. Tits; Kamal Lotfy
The antenna system comprises a reflector which can be motionless, a source device and a device for treating the signals received by the source device. The system comprises means for automatically tracking the satellite that deviates from its initial position to an inclined orbital path.
73	Attitude pointing error correction system and method for geosynchronous satellites	EP89124051.7	1989-12-28	EP0434861B1	1995-02-22	Agrawal, Brij Nandan; Madon, Pierre J.

74	PROCEDE DE RECTIFICATION EN TEMPS REEL D'IMAGES DE SATELLITES METEOROLOGIQUES GEOSTATIONNAIRES	EP90916761.0	1990-10-31	EP0452461A1	1991-10-23	DE WAARD, Johannes; ADAMSON, Jan; BOS, Albert, Martinus
Procédé de rectification en temps réel de données d'images détectées à distance par un instrument de mesure de radiations tel qu'utilisé à bord d'un satellite geostationnaire stabilisé en rotation. Il comprend (1) la prédiction des paramètres de rectification par analyse statistique à partir d'un certain nombre d'images préalables, (2) la lecture d'un nombre limité de lignes d'images brutes pour faciliter la position précise de l'horizon Sud ou de l'horizon Nord, (3) la détermination de la vitesse de rotation du satellite à partir d'un nombre limité de lignes d'images brutes, (4) la redéfinition de l'attitude et des paramètres de pointage del'instrument de mesure en utilisant l'information d'horizon obtenue en temps réel, (5) le calcul informatique de la position de l'autre horizon et du centre de l'image basé sur l'horizon premièrement détecté, (6) la création d'une matrice de déformation faisant suite immédiatement à la détermination du premier horizon, et (7) la rectification en temps réel des données de l'image reçue immédiatement après.
75	Radio frequency broadcasting systems and methods using two low-cost geosynchronous satellites	US48663	1993-04-16	US5319673A	1994-06-07	Robert D. Briskman
High quality audio broadcasts at radio frequencies to mobile receivers at or near the earth's surface are provided by substantially simultaneous transmission of the same signal from two geosynchronous, spatially-separated satellites on the geosynchronous orbit which virtually eliminates multipath fading and foliage attenuation and thereby permits the use of a low-cost space segment.
76	Free return lunar flyby transfer method for geosynchronous satellites havint multiple perilune stages	US79899	1998-05-15	US6149103A	2000-11-21	Jeremiah O. Salvatore; Cesar A. Ocampo
A method is provided for using at least two lunar flyby maneuvers to transfer a satellite from a quasi-geosynchronous transfer orbit having a high inclination to a final geosynchronous orbit having a low inclination. The invention may be used to take the inclination of a final geosynchronous orbit of a satellite to zero, through the use of a first leading-edge lunar flyby and subsequent successive leading or trailing edge lunar flybys resulting in a geostationary orbit.
77	静止衛星位置検出装置、携帯端末装置及びプログラム	JP2015007517	2015-01-19	JP5957101B2	2016-07-27	深町光宏; 山下忠徳

78	Système de tuyères et procédé pour le contrôle d'orbite et d'attitude pour satellite géostationnaire	EP14194874.5	2014-11-26	EP2878539A1	2015-06-03	Amalric, Joël
L'invention propose un système de tuyères (100) pour un satellite destiné à être stabilisé en autorotation sur une orbite géostationnaire, ledit satellite comportant trois axes de référence X, Y et Z, l'axe Y représentant l'axe Nord/Sud et l'axe Z correspondant à une direction de pointage terre. Le système de tuyères comprend un premier ensemble de tuyères (101) configuré pour réaliser le maintien en poste du satellite, le premier ensemble comprenant un nombre pair de tuyères à propulsion électrique, à orientation préréglée, le nombre pair étant au moins égal à 4, lesdites tuyères étant orientées selon trois composantes spatiales, et ayant deux à deux des signes de composantes X et Y différentes.
79	REFLECTEUR OPTIQUE DE SURFACE, POUR UN ENGIN SPATIAL TEL QU'UN SATELLITE GEOSTATIONNAIRE	EP02783228.6	2002-09-30	EP1436197B1	2006-06-14	MAUREL, Gilles
The invention concerns a surface optical reflector, for space craft such as a geostationary satellite. The reflector outer surface (10) comprises a plurality of facets (14) inclined relative to one another and relative to the reflector inner surface (12). Advantageously, the facets (14) form a square-base pyramid whereof the vertex angle is equal to 90 DEG . However, other arrangements are possible within the framework of the invention. The invention enables to enhance the radiative capacity without significantly increasing the volume and the space requirement.
80	REFLECTEUR OPTIQUE DE SURFACE, POUR UN ENGIN SPATIAL TEL QU'UN SATELLITE GEOSTATIONNAIRE	EP02783228.6	2002-09-30	EP1436197A1	2004-07-14	MAUREL, Gilles
The invention concerns a surface optical reflector, for space craft such as a geostationary satellite. The reflector outer surface (10) comprises a plurality of facets (14) inclined relative to one another and relative to the reflector inner surface (12). Advantageously, the facets (14) form a square-base pyramid whereof the vertex angle is equal to 90°. However, other arrangements are possible within the framework of the invention. The invention enables to enhance the radiative capacity without significantly increasing the volume and the space requirement.

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