The principal elements of any rocket-powered flight vehicle are the rocket engine, to provide the propulsive force; the propellants consumed in the rocket engine; the airframe, to contain the propellants and to carry the structural loads; and the payload including any special equipment such as guidance or communication devices.
The rocket engine provides the propulsive forces to accelerate the vehicle by ejecting hot gaseous material at very high speeds through a nozzle. The initial source of the ejected material is the propellant carried in the vehicle in either liquid or solid form. The propellant may be converted to hot gas for ejection by one of a number of possible heating processes in the engine, such as chemical combustion, nuclear fission, etc.
Vertical takeoff from the Earth requires a thrust force that exceeds the weight of the complete missile by some 30 to 50 percent (a thrust-to-weight ratio of 1.3 to 1.5). For easiest engine operation the thrust produced during the entire propulsive period is usually constant, causing the vehicle to be accelerated at a progressively higher rate as the vehicle weight diminishes due to propellant consumption.
In rocket vehicles intended to reach velocities of interest in astronautics, the largest fraction of the missile weight is devoted to the propellants, and the largest volume to the storage of these propellants.
The propellant tanks, and the supporting structure which carries the structural loads imposed during flight and ground handling, comprise the airframe of the flight vehicle. The material in the airframe is considered "dead weight," since it does not contribute directly to the production of thrust or to the useful payload. Rather, the dead weight imposes a limitation on the maximum velocity that a given rocket can achieve-even with no payload.
Another factor contributing significantly to the total dead weight of a vehicle and restricting its maximum performance is any unused propellant trapped in the propulsion system (rocket engine, plumbing, and tanks) at thrust cutoff. In liquid-propellant rockets, two propellant fluids are stored in separate tanks which should be emptied at very nearly the same instant.1 The engine will stop when either propellant is exhausted, and the remaining portion of the other propellant will be trapped as residual dead weight.
The flight velocities required for astronautics far exceed those obtainable with a single rocket unit using conventional propulsion techniques regardless of the size of the rocket. The multistage rocket can provide adequate velocities, however. On this type of vehicle, one rocket (or more) is carried to high speed by another rocket, to be
1 Greenwood, T. L., A High Accuracy Liquid Level Measuring System, Army Ballistic Missile Agency, Redstone Arsenal, Ala., Rept. DTI-TR-1-58, July 29, 1958.
launched independently when the first rocket is exhausted. If, for example, the first stage reaches a terminal velocity of 10,000 feet per second and launches a second stage also capable of developing 10,000 feet per second, the net terminal velocity of the second stage will be 20,000 feet per second.
Staging can be extended to include 3, 4, and more stages to develop higher velocities. The total velocity that may be attained is the sum of the individual contributions of each stage. A practical difficulty will generally restrict the number of stages that can be profitably employed, since the weight of structure required to connect the stages tends to increase dead weight and defeat the purpose of staging.
Control of the flight path of a rocket-propelled vehicle is achieved by altering the direction of engine thrust by one of a variety of methods, including swiveling the engine itself. Velocity control is provided through termination of all rocket thrust at the exact time the desired velocity is reached.
The performance of a rocket is determined largely by the rocket propellant combination and the total amount of usable propellants.2 The performance of propellants is characterized by the specific impulse, a measure of the number of pounds of thrust produced per pound of propellant consumed per second. The unit of specific impulse is lb/lb/sec., or, more simply, seconds. The velocity that a vehicle can attain is directly proportional to the specific impulse of its propellants, all other things being equal. For example, if a given rocket reaches a velocity of 10,000 feet per second with propellants giving a specific impulse of 250 seconds (a typical current value), an increase of 10 percent in specific impulse to 275 seconds would increase the attainable velocity to 11,000 feet per second.
The performance dependence upon the propellant fraction-the fraction of the total vehicle weight accounted for by usable propellants-and the interdependence of specific impulse and propellant fraction are more complex relationships. The maximum velocity increases rapidly as the propellant fraction is made larger, as shown in figure 1, For a given velocity to be achieved, the required propellant fraction is greatly affected by the performance of the propellant combination, as indicated by the three values of specific impulse given in this figure, (The relationship between the propellant fraction and the mass ratio the ratio of takeoff weight to weight at propellant exhaustion, is given by the lower scale of figure 1.) The very high propellant fractions associated with high velocities can only be achieved by severely reducing to a minimum all components that contribute to the weight at propellant exhaustion, including the payload.
2 A more complete discussion of propellants appears In a later section.
The gross weight at takeoff of a rocket vehicle to propel a given payload to a given flight velocity is determined, then, by propellant performance, the minimum practical fraction of dead weight, the number of stages, and the size of the payload to be carried,
Improved materials, design techniques and component miniaturization have led to large reductions in the fraction of dead weight associated with large rockets, The progressively improving state of
the art is graphically illustrated in figure 2, from a value of 0.25 for the German V-2 development in 1939,3 to 0.21-0.16 for the United States Viking rockets in 1949-51,4 to a reported value near 0.08 0.06 for current large rocket developments in the United States.5
The gross weight (and thrust) required for a specified flight objective is very nearly proportional to the size of the payload to be carried. To double the payload, the gross weight-and the thrust- must also be about doubled for similar designs and vehicle configurations. This is the predominant reason for development of large-thrust engines. Additional stages and improved propellants may be used to obtain increased performance from existing rocket booster components.
3 The Missile A-4 Series B as of January 2,1945, translation General Electric DF-71369.
4 Viking Design Summary RTV-N-12a, Glenn L. Martin Co., ER-6534, August 1955.
5 Astronautics and Space Exploration hearings before the Select Committee on Astronautics and Space Exploration, 85th Cong., 2nd sess., on H. R. 11881, April 15 through May 12, 1958.