Method of and device for determining the characteristics and flux distribution of micrometeorites

A micrometeorite impact sensing method of and device for determining the characteristics and flux distribution of micrometeorites are discussed. The method consists of exposing to the micrometeorite environment, a panel of sheet material of a thickness to be punctured by impacting micrometeorites and then scanning the panel with a scanner which produces an output representing the number and size of the puncture holes in the panel. After, exposure, the panel is scanned for puncture holes by illuminating one side of the panel and retracting the panel into its stowage container past a photoelectric scanner which produces an output representing the incident light

A Small, Primary Solar-Electric Propulsion Demonstration Satellite

Interest in the application of primary solar-electric propulsion to high-energy Earth orbital and deep-space missions has increased in recent years because of its inherent propellant economy and the promise of greatly enhanced payload weight capacity for a given vehicle launch weight. This paper describes a small, low-cost electric propulsion demonstration satellite, capable of ascending from low altitude to geosynchronous orbit in ten months or less, using only about 1 kW of propulsive power. Based on available technology of ion thrusters and lightweight solar arrays, and using elements of current light satellite design, this mission is of timely interest as a step in the evolution of future operational electric propulsion transfer vehicles. Generic data on system and mission design for this demonstration vehicle are presented along with relevant literature references

Design of the materials experiment carrier for on-orbit servicing

This paper discussed the MEC system and its mission from the viewpoint of orbit servicing. Information is provided on MEC system requirements, design for on orbit servicing, on orbit servicing operations and rationale and servicing costs

Space station automation study-satellite servicing, volume 2

Technology requirements for automated satellite servicing operations aboard the NASA space station were studied. The three major tasks addressed: (1) servicing requirements (satellite and space station elements) and the role of automation; (2) assessment of automation technology; and (3) conceptual design of servicing facilities on the space station. It is found that many servicing functions cloud benefit from automation support; and the certain research and development activities on automation technologies for servicing should start as soon as possible. Also, some advanced automation developments for orbital servicing could be effectively applied to U.S. industrial ground based operations

Integrated propulsion for near-Earth space missions. Volume 1: Executive summary

Tradeoffs between electric propulsion system mass ratio and transfer time from LEO to GEO were conducted parametrically for various thruster efficiency, specific impulse, and other propulsion parameters. A computer model was developed for performing orbit transfer calculations which included the effects of aerodynamic drag, radiation degradation, and occultation. The tradeoff results showed that thruster technology areas for integrated propulsion should be directed towards improving primary thruster efficiency in the range from 1500 to 2500 seconds, and be continued towards reducing specific mass. Comparison of auxiliary propulsion systems showed large total propellant mass savings with integrated electric auxiliary propulsion. Stationkeeping is the most demanding on orbit propulsion requirement. At area densities above 0.5 sq m/kg, East-West stationkeeping requirements from solar pressure exceed North-South stationkeeping requirements from gravitational forces. A solar array pointing strategy was developed to minimize the effects of atmospheric drag at low altitude, enabling electric propulsion to initiate orbit transfer at Shuttle's maximum cargo carrying altitude. Gravity gradient torques are used during ascent to sustain the spacecraft roll motion required for optimum solar array illumination. A near optimum cover glass thickness of 6 mils was established for LEO to GEO transfer

Small Orbit Transfer Vehicle (OTV) for On-Orbit Satellite Servicing and Resupply

The field of on-orbit servicing of space systems has been studied extensively, and techniques for performing satellite resupply and repair functions have been developed in detail. They are covered extensively in the literature. Based on this background, Microcosm has performed design studies, partly under NASA/MSFC contract, of a small-size, 300 kg-class multi-function Orbital Transfer Vehicle (OTV) that can provide servicing and resupply functions for the International Space Station (ISS). It carries the required payload from a launch vehicle upper stage to the Station, and after berthing it supports servicing activities of the ISS crew members. The vehicle has a payload-carrying capability of 350 kg. The current design includes grappling fixtures specifically designed for ISS berthing which can be eliminated for servicing other satellites. The very strict safety requirements involving ISS access were taken into account in the servicing vehicle design. Repeated ISS servicing sorties to be performed by the OTV are of particular interest, to meet tight revisiting schedules. Extended reuse of the same OTV, once in orbit, allows substantial launch and operational cost savings. Propellant requirements for the servicing sorties are very modest, allowing an extended on-orbit life of this vehicle, with at least 3, but more likely 6 to 8 ISS revisits. The OTV discussed here can be utilized for low-cost servicing of other spacecraft as well. The paper discusses the vehicle’s maneuver sequences and propellant requirements, and describes its design features and its interactions with the ISS. The OTV’s total recurring cost is estimated at less than $35 Million. It would nominally be carried by a light-lift launcher, such as Microcosm’s planned Sprite vehicle, at a projected cost of the order of$2.5 Million

Integrated propulsion for near-Earth space missions. Volume 2: Technical

The calculation approach is described for parametric analysis of candidate electric propulsion systems employed in LEO to GEO missions. Occultation relations, atmospheric density effects, and natural radiation effects are presented. A solar cell cover glass tradeoff is performed to determine optimum glass thickness. Solar array and spacecraft pointing strategies are described for low altitude flight and for optimum array illumination during ascent. Mass ratio tradeoffs versus transfer time provide direction for thruster technology improvements. Integrated electric propulsion analysis is performed for orbit boosting, inclination change, attitude control, stationkeeping, repositioning, and disposal functions as well as power sharing with payload on orbit. Comparison with chemical auxiliary propulsion is made to quantify the advantages of integrated propulsion in terms of weight savings and concomittant launch cost savings

Low-Cost, Minimum-Size Satellites for Demonstration of Formation Flying Modes at Small, Kilometer-Size Distances

In-flight demonstration of close-range formation flying modes is discussed with emphasis on low cost, based on spacecraft design simplicity and minimum size, and hence minimum launch cost. Piggy-back launch of three identical demonstration spacecraft is the preferred option. This demonstration includes comprehensive collection and analysis, on the ground, of relative motion data derived from GPS signals that are received by the participating satellites. In addition to realistically reflecting the known Keplerian and non-Keplerian characteristics as part of the flight dynamics analysis, an important concern is to demonstrate the less well-known effects of differential drag perturbations. These can be deliberately induced by small aerodynamic cross section variations on the satellites. Results of the demonstration mission will benefit currently projected and other future missions that require close-formation flying, e.g., radar and visual observation missions, by proving feasibility, assisting detailed and comprehensive operations planning, and helping diminish their potential risks

Responsive, Low-Cost Access to Space with ELVIS — An Expendable Launch Vehicle with Integrated Spacecraft

The ELVIS (Expendable Launch Vehicle with Integrated Spacecraft) concept involves: (1) dropping off the upper stage of the launch vehicle as low as possible, with integral low-thrust propulsion taking the spacecraft to its final orbital destination; (2) using the spacecraft bus to provide the avionics functions needed to fly a launch vehicle to orbit so as to avoid the duplication of avionics hardware and software between the satellite bus and the launch vehicle. The result is a reduction in the parts count, weight, and cost of the launch vehicle. There are major benefits associated with early staging — the upper stage can reenter safely without a retro burn, and the mass-to-orbit available from small launch vehicles is significantly increased. The mass gain will depend on the hardware configuration and the orbit destination, but can be as much as a factor of two or more for some low Earth orbits. In addition, the spacecraft bus operates from the time of launch and can begin the mission essentially as soon as the spacecraft reaches its operational orbit or, in some cases, even before. The small spacecraft thus achieves a new level of responsiveness, allowing spacecraft to be launched in response to rapidly changing circumstances. This paper describes a representative ELVIS configuration and performance gains for typical mission destinations, and sample applications that are enabled or made more efficient by the use of this approach. Technical issues and tradeoffs associated with this design will be discussed