The Stennis Space Center acts as the program office for remote sensing activities on behalf of the NASA Office of Commercial Programs. Stennis is supplying systems engineering support to determine the technical and economic feasibility of a commercial remote sensing small satellite. Study efforts focus on the systems design of launch vehicle, sensor system, communications, attitude control, power, computational processing, and ground station capabilities. Results reveal that an example small satellite system, with a limited capability remote sensing payload, could be deployed for less than $20M. Technological advances and overall cost reductions in space systems may enable industry to develop new and expanded information service markets.
On January 21, 1990, Weber State University had its second small satellite called WEBERSAT launched into an 800 Km polar orbit on the Arianespace Ariane V35 flight from Kourou, French Guiana. This launch culminated nearly two years work by undergraduate engineering technology students and faculty at Weber State University, and local volunteer engineers. The successful development of this satellite was the result of joint efforts with education and industry. This project is demonstrating that a low cost spacecraft can use inexpensive sensors to provide useful space data. This paper is an overall view of the satellite and other papers in the proceedings will provide more details on operational performance.
Three experimental small satellites were developed and flown by ISRO for imaging purposes. This paper gives the details of the payloads of these satellites. The first satellite was ROHINI-D1 which carried a single-band panchromatic band camera. The ROHINI-D2 satellite had a two-band camera using linear detector arrays and had a capability to identify features using the ratio of signals in two spectral bands. The SROSS-2 satellite carried a stereo imaging camera with capabilities of getting triple images along the ground track.
The paper examines the application of small satellites for water cycle experiments, giving attention to alternative satellite configurations, subsystems requirements, attitude control requirements, and launch strategies. Particular emphasis is placed on the attitude and orbit control, which will be low-cost while providing the high attitude precision required by earth-observing sensors. The LEOSTAR configuration for water cycle experiments is considered.
The 'Eagle' Class small satellite is a modular Lightsat capable of supporting a wide variety of LEO applications. This paper highlights the key parameters of the spacecraft - power, stabilization, payload accommodation, etc., which make it useful as a platform for scientific experiments. The satellite consists of modules such as core, payload, deployables and orbit insertion subsystem which are integrated into a stacked configuration. The 38 in. diameter is sized for the Pegasus or Air Force SLV. Although considered a 'Lightsat', the satellite can offer as much as 500 watts of power and .01 degree attitude knowledge. It is suitable for a wide variety of earth observation and sensing missions. The satellite can be commanded by an economical desk-size Master Ground Station.
Athena is a low cost imaging satellite system capable of providing high resolution earth imaging in a single band between .5 and .73 microns. The Athena payload consists of two optically fast cameras, coupled with linear charge coupled devices to provide 5 meter resolution over a 700 km swath from a 700 km polar orbit. A patented chip allows implementation of vector quantization image compression technology on board the spacecraft. This technology provides up to 12 to 1 image compression, dramatically reducing satellite to ground datalink bandwidth, on-board storage, and spacecraft power requirements while returning 'statistically lossless' images to the ground. The satellite and payload design contain few moving parts, thereby achieving high reliability and low cost, and requiring minimal operational support. The ground station design permits fixed or mobile operation and can be supported by a single technician.
The T-SAT design was developed to meet the need for multifunctional low-cost small satellites for carrying scientific and/or commercial payloads. The basic designs are rugged, simple, and mass-producible, using both commercial and aerospace technology. An overview of the three-axis stabilized T-SATs for use in very low earth orbits, low earth orbits, and geostationary orbits and beyond is presented.
The objective of the TECHSTARS initiative is to examine the architectures associated with small space systems and to incorporate the advanced technologies required for the 1990's. The TECHSTARS program comprises a line of small satellites and ground systems for scientific, commercial, civil, and military applications. The TECHSTARS satellites weigh from 100 to 500 pounds and have significant onboard processing capabilities. The spacecraft, the ground systems, and technology development in the TECHSTARS program are addressed here.
The astronaut deployable satellite (ADSAT) is the third in a series of satellites being designed and built by students, faculty, and local engineers at Weber State University. These satellites have been an excellent vehicle for providing realistic engineering experiences for the students. The ADSAT is designed to be launched from the space shuttle by an astronaut. It will be capable of sending back to earth voice messages concerning the onboard experiments. The synthesized voice messages can be received by inexpensive scanners on the ground.
This paper describes two complimentary experiments, the purpose of which is to measure the forces involved in the deployment in Earth orbit of a tether and its end mass. These two experiments are called TAS, (Three-axis Accelerometer System), and SEASIS, (SEDS Earth Atmospheric and Space Imaging System). The two experiments are flown as a student experiment package within a small satellite. This satellite (SEDSAT 1), which constitutes the end mass, will be deployed propulsively into a higher, permanent orbit utilizing a tethered system. These experiments are proposed to fly on a McDonnell Douglas Delta II as a secondary payload on a Global Positioning System (GPS) mission. This flight will be the third demonstration flight of the Small Expendable Deployer System (SEDS), a project managed by the NASA Marshall Space Flight Center, Huntsville, Alabama.
A series of low-cost spacecraft buses have been designed which are sized to host dedicated sensor packages to be launched into low earth orbit on small inexpensive launch vehicles or in a Get-Away-Special Canister in the Space Shuttle. These buses facilitate realizing rapid on-orbit status by employing a standard system module with flexible options for attitude control and power. The buses may be configured for operational or experimental missions with lifetimes of a few months to several years. The series consists of three classes of spacecraft, each one sized to satisfy a range of requirements for mission in the areas of remote sensing, space experiments, technology testing, microgravity studies, surveillance, environmental and geophysical measurements, location determination, real-time and store-and-forward communications, and other space applications. Integral to each class of bus is a communications and data handling package which provides programmable mission activities under operator control through a companion low cost ground station. Parameters for size, power, and weight are scaled to the requirements of the payload. These buses permit rapid implementation of operational or experimental remote sensing programs within limiting monetary constraints.
W.J. Schafer Associates Inc. has conducted an assessment of the capabilities of small satellites to fly material experiments in low earth orbit in order to measure the environmentally induced material degradation. A potential experiment was chosen for the assessment which requires spacecraft pointing in the forward velocity vector (ram direction), a minimum of power, and low data storage. A satellite designed to be launched from a Space Shuttle Get Away Special (GAS) is nominally able to meet these requirements if a simple drag stabilization system is used, if all the available surface area is used for solar panels and if the experiment's power requirement is minimized through the use of passive thermal control.
Pegasus and the PegaStar integrated spacecraft bus are described, and an overview of integration and launch operations is provided. Payload design issues include payload volume and mass capability, payload interfaces, and design loads. Vehicle and payload processing issues include integration and handling methods, facilities, contamination control, and launch operations. It is noted that Pegasus provides small satellite users with a cost-effective means for delivering payloads into the specific orbits at the optimal time to meet the most demanding mission requirements. PegaStar provides a flexible cost-effective means for providing long-term on-orbit support while minimizing total program risk and cost.
The RESERVES program sensor concept, a multiwavelength optical system with two standard focal planes, handles many missions by using commandable readout and processing options. Performance (coverage rate, resolution, spectral data) and rapid operational readiness are achieved with the size, power, and communication link constraints of a 450-kg spacecraft. Missions achievable include weather and oceanographic mapping, earth-disk scanning for strategic warming, focused rapid-repeat theater scanning for intratheater missiles, land and ocean remote sensing, and space-object surveillance.
After reviewing DARPA's Advanced Space Technology Program (ASTP) and discussing the need for the ASTP, the paper examines DARPA space technology initiatives. Space technology initiatives include the UHF, SHF, and EHF initiatives, laser communications, optics stabilization, and optics and lasers. Supporting technology initiatives include initiatives for onboard computing; mechanical components; autonomous guidance, navigation, and control; and propulsion, power, and thermal.
Advanced electroopical sensors and surveillance system technologies are being developed that would lead to future space-based surveillance systems. These multispectral sensors would operate in the visible and MWIR wavebands and be configured to be compact, lightweight, and low power to permit deployment on lightsat-type platforms and launch vehicles. The sensors for multi-spectral space surveillance are small electrooptical systems which use large staring focal planes to provide maximum detection capability through integration. Operation in visible and MWIR bands permit surveillance of satellites in sunlight and in earth shadow. Four to six small satellites with these sensors at low altitude provide essentially instantaneous coverage of orbital space out to geosynchronous altitudes. A baseline sensor design has been derived along with a total sensor payload configuration with size, weight, shape and power requirements compatible with those for a satellite deployed by the Pegasus launch system.
This paper describes a new approach to location determination and remote sensing data collection, utilizing a satellite configuration that supports a unique antenna payload designed to scan the surface of the earth to receive location and messaging data from inexpensive end-user transceivers and relay this information to ground centers for processing and distribution.
The Earth's magnetic field can provide attitude reference for satellites. Because of the nature of the terrestrial magnetic field, magnetic sensing is primarily used to sense azimuth, with a horizon sensor providing reference in the other two axes. The magnetometer azimuth sensor is small, light, and inexpensive. A new solid-state magnetometer based on magnetoresistive techniques can provide azimuth determination with an accuracy equal to existing techniques. It has even more favorable advantages, however, in size, weight, power consumption, and cost compared to conventional techniques. A magnetic azimuth sensing system based on this magnetometer should provide azimuth information to about 0.5 degrees, a limit imposed by uncertainties in the Earth's field rather than limitation in sensitivity of the solid-state sensor.
A combined absolute/relative navigation experiment, currently under analysis in Italy, is described. Two space vehicles are constrained to take measurements from the same GPS satellites. The absolute and relative state vectors are determined by a Kalman filter that processes the differences between the two vehicle measurements relevant to the same GPS satellite and the absolute measurements of each vehicle. The experiment is implemented using two small coorbiting satellites, Pegasus/Scout II compatibles, flying in circular orbits at an altitude of about 400 km and manuevering to vary the distance between the two satellites. Absolute and relative position and velocity are autonomously computed onboard the vehicles with accuracies sufficiently high to allow GPS-based rendezvous operations and autonomous navigation.
The Small Expendable Deployer System developed by Langley Research Center has an opportunity to deploy a student experiment package in mid-summer 1993. This paper outlines a preliminary bulk thermal analysis of this small satellite. The satellite thermal design is passive in elliptical low earth orbit. Calculations of the equilibrium space craft temperature uses a total energy balance method. Transient response during eclipse periods is also addressed.
An optical computer has been constructed employing an artificial intelligence, early vision, processing technique to perform precision horizon location in scanning horizon sensors. It is well known that the second derivative of the horizon signal crosses zero at the point of maximum slope and is thus an optimum point to minimize instrument errors. In contrast to conventional techniques the Bulls EyeTM Locator performs the double differentiation optically, thus circumventing most of the noise limitations imposed by electronic differentiation and analog/digital signal processing. Differentiation is implemented spatially by subtracting the signals from an inner circular and an outer annular field of view (FOV). The result is a method of horizon location which is insensitive to scan speed and crossing angle variations, along with inherently rejecting radiance errors. Optimization of on-orbit performance is reviewed in detail. Several embodiments of the locator are discussed, including a proprietary single-element pyroelectric detector. Technical implications are explored, specifically, our development of an extremely low cost, high accuracy, wide dynamic range family of horizon sensing instruments.
A miniaturized, low-power parallel processor for space applications is under development by Space Computer Corporation for DARPA's Advanced Space Technology Program. The basic goal of this project is the reduction, by an order of magnitude or more, of on-board processor weight, size, and power consumption for space-based sensor systems. The approach described here for achieving this goal is to use low-power VLSI devices which maximize throughput per watt, together with three-dimensional hybrid wafer-scale integration and packaging technology. In its prototype version, a 12-node processor will have a peak throughput greater than 1.2 GFLOPS and occupy a volume less than 15 cubic inches.
Nickel-Hydrogen (Ni-H2) batteries have become the technology of choice for both commercial and defense-related satellites in geosynchronous orbits. Their use for low-earth-orbit (LEO) applications is not as advanced, but seems just as inevitable because of their inherent advantages over nickel-cadmium batteries. These include superior energy density, longer cycle life, and better tolerance to over-charge and reversal. Ni-H, cells have the added advantage in both construction and operation of not presenting the environmental possibility of cadmium pollution. Unfortunately, but necessarily, the design of these cells has been driven to high cost by the sophistication of the satellites and their uses. Now, using most of the same concepts but less costly materials and techniques, a low cost, small cell design has been developed. Combined with the concept of the common pressure vessel, this new design promises to he ideal for the "small-sat" and commercial markets which, increasingly, are calling for large numbers of less-expensive satellites.
Webersat, launched January 21, 1990 into a 800 km polar orbit, has a micrometeorite impact detector plate on one face. This paper describes the results of measurements made with this detector since launch and the correlation of this data with satellite altitude, location and meteorite showers.
Webersat has exhibited strange dynamic behavior as confirmed by telemetry; the spin rate has decreased from 5 to 0.05 rpm over the last year. The elongated shape of Webersat is noted to result in an unstable spin about the Z-axis. It is proposed that, in the future, spin-stabilized satellites should be designed to be short and fat about their spin axis rather than tall and thin. The Webersat space frame can be made to spin about the Z-axis if the masses are redistributed so as to produce an overall increase in the moment of inertia about the spin axis.
An important addition to the emerging commercial space sector is Standard Space Platforms Corporation's comprehensive low-cost flight service delivery system for small and developmental payloads. Standard provides a privately funded, proprietary, value-added transportation service which dramatically reduces cost and program duration for compliant payloads. It also provides a business-to-business service which is compatible with business investment decision timing and technology development cycles.
An architecture is proposed which offers a structure for many different small satellite types to be integrated into a system constellation to meet the multiple demands of today's various users. An approach to evolving such a constellation is proposed that supports operations for weather prediction mapping, communication, and reconnaissance. Alternative designs for supporting sensor viewing and communication are discussed. Alternatives are offered for multimission networking and connectivity to users, and attention is given to options for controlling and operating a constellation of tens or hundreds of satellites intermixed with different types of capabilities and designs.
This paper describes the spectrum allocation and licensing issues that must be faced by any company contemplating development of a private remote-sensing satellite system. Pursuant to the Land Remote-Sensing Commercialization Act of 1984, private remote-sensing satellite systems must be licensed by the Secretary of Commerce. Frequency assignment and licensing by the Federal Communications Commission is also required. Although the legal framework is in place, there are still issues to be resolved in the context of specific applications, including those issues arising from the inherent tension between commercial remote-sensing activity and national security concerns.