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A high Mach number, high altitude aircraft, utilized as a launch platform for a reusable air-launched rocket, can provide several benefits for access to low Earth orbit. Performance figures are presented for reusable, winged rocket stages launched from a large supersonic aircraft comparable to the large XB-70 aircraft. The XB-70 achieved mach 3.1, 73,000 ft altitude flight with conventional turbojet propulsion in 1964. Advantages of air-launch from this condition include a 23,644 ft/s ascent-to-orbit ideal delta-velocity (a 22% reduction compared to a ground launch), reduced drag incurred by rocket flight beginning at an atmospheric density one-twentieth that of sea level, and the capability to launch at any latitude. A reusable rocket of 343,000 lbs launched from an appropriated sized aircraft could deliver a 23,500 lb externally mounted payload pod to low Earth orbit utilizing a 523 sec Isp advanced rocket engine. Fluorine/lithiumhydrogen engines achieved 523 sec in USAF and NASA development programs of the 1960’s and 1970’s, an increase of 70 sec Isp over that of the O2/H2 SSME. Together, the 22% delta-V and 70 sec Isp advantages would enable significant reductions in rocket weight compared to a ground-launched all rocket vehicle delivering the same payload to low Earth orbit. Also presented are data for a similarly configured 465 sec Isp O2/H2 air-launch rocket. Fluorine reactivity pasivation and engine development history is also discussed.

Systems analyses were performed for a Titan Explorer Mission characterized by Earth‐Saturn transfer stages using solar electric power generation and propulsion systems for primary interplanetary propulsion, as well as chemical propulsion for capture at Titan. An examination of a range of system factors was performed to determine their effect on the payload delivery capability to Titan. The effect of varying launch vehicle type, solar array power level, ion thruster number, specific impulse, trip time, and Titan capture stage chemical propellant choice was investigated. The major purpose was to demonstrate the efficacy of applying advanced ion propulsion system technologies like NASA’s evolutionary xenon thruster (NEXT), coupled with state-of-the-art (SOA) and advanced chemical technologies to a Flagship class mission. This study demonstrated that a NASA design reference mission (DRM) payload of 406 kg could be successfully delivered to Titan using the baseline advanced ion propulsion system in conjunction with SOA chemical propulsion for Titan capture. In addition, the solar electric propulsion system (SEPS)/chemical system of this study is compared to an all-chemical NASA DRM. Results showed that the NEXT-based SEPS/chemical system was able to deliver the required payload to Titan in 5 years less transfer time and on a smaller launch vehicle than the SOA chemical option.