CRYOGENIC-PROPELLANT HEATING IN THE THERMAL ENVIRONMENT OF SPACE

Heat is transferred to a propellant tank in space from the Sun, planets, and on-board components.Radiation and conduction from adjacent components are relatively simple to define and are a familiar problem to the designer of Dewars for ground installations.Unless heat transfer to cryogenic propellants from these on-board sources is extremely small, there will be excessive propellant vaporization.The heat transfer by radiation from the Sun and planets to propellant tanks is discussed in detail as is the effect of using various thermal barriers to reduce propellant heating.An extensive list of equations is included to summarize the results of the analytical derivations for each particular thermal-protection system.Operation in a planet orbit, in general, subjects the propellants to a time-varying radiation environment.The attitude of the tank with respect to both the Sun and a nearby planet must be known at all times in order to estimate propellant heating rates accurately.The choice of orbit altitude provides some possibility for alleviation of adverse propellant heating effects.Of major concern in the preliminary analysis of a space vehicle is the maximization of payload weight.The ultimate effect of the thermal environment of space on the design of a particular vehicle is a weight penalty directly chargeable to this environment.A method of calculating and optimizing this weight penalty is included for a hypothetical Mars trip with a hydrogen-oxygen-fueled chemical stage.