burn rates. Burn rates for operational rocket propellants are typically between 0.3 and 0.8 inches per second, which would generate unacceptably long burn times (~40 minutes for the 110-foot-long solid rocket motor units (SRMUs) used on the Titan IV booster if the propellant burned from bottom-to-top—like, for instance, a cigarette). Large, solid rocket engines have propellant cores cast with convoluted openings to provide radial burning with larger surface areas; this leads to higher gas generation rates and shorter burn times (about 2 minutes for the Titan IV solid boosters). Nanopowder propellants offer much higher burn rates, so simplified core designs with higher average propellant density could be used. Technical issues include determination of nanopowder oxidation rates in contact with the fuel/oxidizer to establish propellant storage lifetimes, control of engine instabilities with faster-burning propellants, and the possible use of coatings on the nanopowders to chemically stabilize them.

Liquid propellant rocket engines offer higher exhaust velocities than solid rockets in return for increased complexity and cost. Higher specific impulse translates into reduced propellant mass requirements for any particular mission. Liquid hydrogen and oxygen provide the highest practical specific impulse (about 450 seconds), while liquid hydrazine (or its derivatives) and nitrogen tetroxide offer noncryogenic storage with hypergolic ignition for increased reliability at reduced specific impulse (about 320 seconds). Payload delivery performance is a function of specific impulse, wet/dry mass fractions, and the number of stages used. Micro- and nanoengineered materials with increased strength-to-weight ratios, when realized, will improve the wet/dry mass fractions and result in more delivered payload per unit launch weight.

A complete liquid propulsion system consists of a thrust chamber and nozzle, propellant piping and controls, and pumps or a high-pressure gas system (to pressurize the propellant tanks), which feeds propellants to the combustion chamber. With current MEMS technology it is now feasible to micromachine all of these components at the millimeter scale. The first MEMS dime-sized, 3-pound-thrust bipropellant rocket engine has been tested. This device, developed at the Massachusetts Institute of Technology with NASA funding, was fabricated out of six bonded, micromachined silicon layers; it uses oxygen and methane as propellants. Owing to its small size and favorable scaling relationships, its thrust-to-weight ratio is far greater than that for traditional bipropellant thrusters. Future engines may be developed at larger or smaller thrusts as the system requirements and technology allow. Another example of a smaller, less ambitious bipropellant thruster is the “microjet” developed in the United Kingdom at the Defence Establishment Research Agency. This 13-millimeter-long, 2-gram-mass thruster produces 63 millinewtons of thrust using hydrogen peroxide and kerosene. Obviously, multiple units can be used in parallel to achieve higher thrusts.