Horizon Trackers for Lunar Guidance and Control Systems

Abstract

That horizon-sensing instruments have a future in space vehicles has already been aptly demonstrated in the TIROS I satellite and rocket tests run as early as the summer of 1959. It is the purpose of this paper to discuss the extension of these techniques to the problem of lunar guidance and control.

In one sense, the lunar problem is somewhat simpler than the problem of the earth horizon-track er problem due to the lack of an atmosphere near the lunar surface which might obscure the planet-space interface. However, since sightings on the earth will probably also be of interest to a lunar guidance and control system, the problem of the earth horizon tracker is also considered in this paper, including the effects of the earth’s atmosphere on the accuracy of the instrument.

A discussion of the oblateness effect, as discussed by Roberson, is briefly reviewed, followed by a detailed discussion of the refraction effects of the earth’s atmosphere; it is shown that under certain circumstances, the two effects cancel each other. For high-accuracy instruments, compensations must be introduced. The problem of cloud cover is briefly discussed, and a possible solution, which might be of particular significance to a vehicle re-entering the earth’s atmosphere for recovery, is offered. The calculations of radiation, necessary for the design of the instrument, are outlined, including absorption by atmospheric constituents, along with some of the theoretical limitations of the instrument.

The various design considerations, including scan mechanization, choice of detector, signal processing, and other factors, are discussed; however, no attempt is made to offer an optimum design, since this will depend strongly upon the specific mission, the vehicle used, and accuracy requirements. It is hoped, however, that enough information is presented to assist the designer of a guidance and control system in his’ choice of a horizon-tracking instrument.

The authors wish to acknowledge the help of Mr. R. Kroeger and Mr. R. Wrobel for their assistance and encouragement during the preparation of this paper, and for their aid in the fulfillment of the contract under which part of the work was accomplished. The work described was performed under Contract No. AF 33(616)-6674 sponsored by Wright Air Development Division USAF under subcontract from Systems Corporations of America, Los Angeles, California.