Raytheon Company

The Boeing Company

Hughes Aircraft Company

Hughes Electronics Corporation

Christopher M Cosner

Joseph Daniel Khair

Michael S Baldwin

John F Stickelmaier

John J Wootan

Stuart F Bockman

Dennis Yoichi Nakasone

Orion Derek Radtke

Pavan Daryanani

David L Behrin

Scott M Matsunami

A space vehicle system includes a first star tracker disposed at a first location on or near the space vehicle, the first star tracker configured to obtain first images of a space object and stars and a second star tracker disposed at a second location on or near the space vehicle, at a distance D from the first location. The second star tracker is configured to obtain second images of the space object and the stars and the first images and the second images being stereoscopic images. The system also includes a processor configured to determine an estimate of a range from the space vehicle to the space object based on the first images and the second images.

Star trackers for range determination in rendezvous and proximity operations

Systems and methods are provided for processing and storing remote sensing data. An exemplary method includes collecting information regarding a target based on an objective; receiving the information regarding the target; temporally storing the information regarding the target; deriving, using the information regarding the target, a first data fingerprint based on the objective; providing a library configured to store at least one data fingerprint; determining whether the library includes a second data fingerprint having the same objective; upon determining that the library includes the second data fingerprint, determining whether there is sufficient difference between the first data fingerprint and the second data fingerprint; and in response to determining there is sufficient difference between the first data fingerprint and the second data fingerprint, storing the information regarding the target in a remote sensing platform.

System and method for remote sensing using data fingerprinting methodologies

An electrical power system has a dual battery configuration that enables sufficient power supply for a spacecraft bus and a payload module being carried by the spacecraft. During a sunlight power mode, power is drawn from a solar array of the bus to power a low-discharge payload of the spacecraft and a high-discharge payload of a payload module. During the sunlight power mode, a low rate discharge battery and a high rate discharge battery are charged by a battery charge management unit of the spacecraft bus. During an eclipse power mode, the low rate discharge battery powers the low-discharge payload of the spacecraft and the high rate discharge battery powers the high-discharge payload of the payload module. The high-rate discharge battery may also be used to power the high-rate discharge payload in the sunlight power mode to meet its high current demands to meet a flexible mission operations.

Modular electrical power subsystem architecture

A self-contained payload module and method of deployment of a payload includes a housing that is configured to engage a propulsive payload adapter hub, at least one deployable payload that is arranged in an interior cavity of the housing and deployable using the propulsive payload adaptor hub, a payload electronics system arranged in the interior cavity, a thermal control system in communication with the at least one payload and arranged in the housing, and at least one antenna that is arranged on the housing and configured for wideband communication.

Self-contained payload accommodation module

An exemplary satellite includes a pair of payloads coupled to one another by a flexible boom, where the flexible boom is configured to enable easy manual engagement of the pair of payloads with an associated attachment hub, and also to provide a passive release force for deploying at least one of the payloads in a direction outwardly from the attachment hub. Potential energy for enabling the passive release is provided as stored strain energy in the flexible boom when flexed for attachment of the payloads to the attachment hub. The strain energy is released upon release of at least one of the payloads from the attachment hub, which release may be by way of a non-complex, non-exotic attachment mechanism. Additional payloads may be connected in series to the pair of payloads, with a flexible boom connecting adjacent payloads. The additional payloads may be released from the attachment hub via a non-complex or complex attachment mechanism.

Flexible satellite for deployment from attachment hub

A system, method, and apparatus are disclosed for agile dedicated spacecrafts for spinning microwave imagers and sounders. In one or more embodiments, the system, method, and apparatus involve an agile, zero net-momentum, spinning space vehicle, which includes a body and an instrument package. In one or more embodiments, the instrument package is mounted directly onto the body such that the space vehicle can point the instrument package not only along the space vehicle's orbital velocity vector, but in any direction within the space vehicle's field of regard. The space vehicle's spin axis is aligned with the instrument package's scan axis. The space vehicle experiences zero net-momentum on orbit by including a counter-rotating momentum storage device. In one or more embodiments, the instrument package is a scanning microwave imaging/sounding instrument, which is utilized as an Earth climate and weather sensor.

Agile dedicated spacecraft for spinning microwave imagers and sounders

An agile, spinning spacecraft has a first body portion and a second body portion which is rotatably coupled about a first rotational axis to the first body portion. A pair of planar solar cell arrays are rotatably coupled to the second body portion along a second rotational axis that is preferably orthogonal to the first rotational axis. The first body portion can be spun if desired and the first rotational axis can be directed by an attitude control system to have selected orientations. The solar cell arrays are rotated about the first and second rotational axes in accordance with steering laws which maintain the arrays in an orthogonal relationship with a Sun line.

Agile, spinning spacecraft with sun-steerable solar cell array and method

A control system (40) and procedure (50) that uses repetitive control to control repetitive error sources such as result from the phenomenon of thermal shock experienced by a spacecraft (10). Repetitive control is a 'learning algorithm' that substantially eliminates errors in a stable control system (43) that performs highly repeatable tasks. The repetitive control system (40) and procedure (50) employs a repetitive signal generator (41) (G.sub.r), to store a sensed short-term error signal during each cycle and to process it to generate a signal that compensates for the error. The repetitive control system and procedure integrates (52) the error signal over several cycles, multiplies (53) the integrated value by a predetermined gain factor, and sums (54) it with the sensed error signal before the sensed error reaches the stable controller (43). An output filter (42), G.sub.f, which may be a multirate-sampling output filter, may be used to increase stability robustness of the repetitive control system (40) and procedure (50). The present repetitive control system (40) and procedure (50) substantially eliminates pointing error caused by thermal shock disturbance, for example. The repetitive control system (40) and procedure (50) 'learns' the errors that exist in the system and over several cycles, converges on a substantially exact compensation for these errors. Implementation of the present control system (40) and procedure (50) allows the spacecraft (10) to maintain precise pointing throughout an eclipse and thereby allow the spacecraft (10) to perform its mission despite the presence of thermal shock disturbances or errors caused by other repetitive error sources.

Growing community of inventors

Manhattan Beach, CA, United States of America

Christopher M Cosner