the latter is used for the Delta IV RS-68 or the Saturn V upper-stage J-2 engine. An example of the closed-cycle engine is the staged-combustion design used on the space shuttle main engine (SSME), which is in the process of being simplified to a lower cost expendable version called the RS-25E. (The RS-25E is the baseline engine selected by the Air Force for the RBS expendable upper stage.) Yet another closed-cycle LO2/LH2 engine could be based on an expander cycle such as is used for the RL-10 family of upper-stage engines.

While an LO2/LH2 MPS will result in a much higher specific impulse (Isp) (approximately 390 sec at sea level) than a hydrocarbon-fueled engine, it is a nonoptimal choice for the RBS first stage, owing to the extremely low propellant density of hydrogen and its deep cryogenic properties, which must be maintained at approximately −420°F. The use of these deep cryogens results in large, heavy, and more complicated aerodynamic configurations as well as a relatively poor stage mass fraction. Since mass fraction is just as important as Isp in the basic vehicle “rocket equation” design solution, a higher density, more easily storable fuel such as kerosene is a better choice than liquid hydrogen for the RBS booster. The many operational advantages of using kerosene in the first stage, as compared to cryogenic hydrogen, provide a further rationale for the selection of a hydrocarbon fuel.

The selection of a higher density fuel leads to the second set of options for the RBS MPS, which is engines that operate using liquid hydrocarbon fuels, such as rocket propellant (RP)-1, as recommended and advocated by the Air Force in their presentations to the committee. Liquid methane or even ethane or propane (such as the main constituents of liquefied natural gas) might also be a good choice for a higher density, better performing fuel, but there is only limited technology experience with these fuels and LO2 oxidizer in rocket combustion devices and in vehicle flight experience. Thus, while the committee believes that these fuels might be a good choice for future advanced launch systems, rocket engines designed for using liquefied natural gas-type fuels are not currently at a sufficient TRL for serious consideration.

The committee therefore agrees with the Air Force baseline selection of LO2/RP-1 for the RBS MPS. Again there are also two suboptions for this propellant combination: open cycle, either gas generator or tap-off, and closed cycle ORSC. The ORSC engine is physically the highest performance (Isp) design approach for RP-1 fuels for several fundamental reasons, including these: (1) operation at high chamber pressure which provides higher net Isp because of improved combustion efficiency; (2) higher sea level (liftoff conditions) area ratio; and (3) typically much higher engine thrust-to-weight ratio. Key characteristics for open-cycle gas generator or tap-off cycle engines versus ORSC closed-cycle engines are summarized in Table 3.2.

Table 3.3 summarizes the advantages of an ORSC LO2/RP-1 engine over an open-cycle gas generator and lists the issues/concerns as well. The turbine drive gas from the fuel-rich gas generator presents a more benign condition for all the materials in the hot-gas flow path but forces the turbine to run at much higher operating temperatures in order to achieve the necessary turbine drive power.

As discussed previously, the RBS uses the booster main propulsion system to accelerate the upper stage(s) and payload to the staging velocity and to provide sufficient impulse to allow the first stage to glide back to the launch site. With these dual demands on the main propulsion system, the first-stage sizing becomes very sensitive to the specific impulse and thrust-to-weight ratio of the LO2/RP-1 engine. For this reason, the committee agrees with the Air Force assessment that the ORSC cycle is the preferred engine cycle for the RBS system.

A typical ORSC engine is depicted schematically in the very simple diagram in Figure 3.1. Liquid oxygen and fuel are fed into a high-pressure preburner, which initially combusts at greater than stoichiometric mixture ratios (MR = 10-20), and then the initial combustion products are quenched downstream in the preburner chamber with large amounts of LO2. The LO2 is usually injected down below in the preburner chamber through some type of tangential slots, such that the effluent gas leaves the preburner at an MR of 60-80 or so and an exhaust

TABLE 3.2 LO2/RP-1 Typical Range of Operating Conditions

Cycle Nominal Chamber Pressure Range (psia) Vacuum Isp Range (s) Thrust/Weight Ratio
Open, gas generator 500-1,000 300-315 70-80
Closed, oxygen-rich staged >2,000-3,500a 325-350 100-120

a More typical for Russian engines such as RD-170 and NK-33.



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