Space Launch 28 years of studies

A Personal Journey By John W Livingston 2011

1 Space Launch 28 years of studies or Engineering & Management run amuck

A Personal Journey By John W Livingston 2011

…and all I got was this lousy T-shirt

2 1958-64 Early studies (pre-me) Reusable Rockets & Aerospaceplanes • Wide ranging studies • Numerous combined cycle engines investigated

Inadequate technology base, but really hot stuff.

The US Air Force's aerospaceplane project encompassed a variety of projects from 1958 until 1963 to study a fully reusable spaceplane. A variety of designs were studied during the lifetime of the project, including most of the early efforts on liquid air cycle engines (LACE) and even a nuclear-powered ramjet. The effort was started largely due to the work of Weldon Worth at the Wright-Patterson AFB, who published a short work outlining a manned spaceplane. AF officials were interested enough to start SR-89774 (study requirement-) for a reusable spaceplane in 1957. By 1959 this work had resulted in the Recoverable Orbital Launch System, or ROLS, based around a LACE engine, known at the time as a Liquid Air Collection System, or LACES. Further work showed that more performance could be gained by extracting only the oxygen from the liquid air, a system they referred to as Air Collection and Enrichment System, or ACES. A contract to develop an ACES testbed was placed with Marquardt and General Dynamics, with Garrett AiResearch building the heat exchanger for cooling the air. The original ACES design was fairly complex; the air was first liquified in the heat exchanger cooled by liquid hydrogen fuel, then pumped into a low pressure tank for short term storage. From there it was then pumped into a high pressure tank where the oxygen was separated and the rest (mostly nitrogen) was dumped overboard. In late 1960 and early 1961 a 125 N demonstrator engine was being operated for up to five minutes at a time. In early 1960 Air Force offered a development contract to build a spaceplane with a crew of three that could take off from any runway and fly directly into orbit and return. They wanted the design to be in operation in 1970 for a total development cost of only $5 billion. Boeing, Douglas, Convair, Lockheed, Goodyear, North American, and Republic all responded. Most of these designs ignored the ACES system and instead used a scramjet for power. The scramjet had first been outlined at about the same time as the original LACES design in a NASA paper of 1958, and many companies were highly interested in seeing it develop, perhaps none more than Marquardt, whose ramjet business was dwindling with the introduction of newer jet engines and who had already started work on the scramjet. Both Alexander Kartveli and Antonio Ferri were proponents of the scramjet approach. Ferri successfully demonstrated a scramjet producing net thrust in November 1964, eventually producing 517 lbf, about 80% of his goal. Later that year a review suggested that the basic concepts of the aerospaceplane were far too new for development of an operational system to begin. They pointed out that far too much was being spent on development of the aircraft, and not nearly enough on basic research. Moreover, the designs were all extremely sensitive to weight, and any increase (and there always is one) could result in all of the designs not working. In 1963 the Air Force changed their priorities in SR-651, and focused entirely on development of a variety of high- speed engines. Included were LACES and ACES engines, as well scramjets, turboramjets and a "normal" (subsonic combustion) ramjet with an intake suitable for use up to Mach 8. In October a further review concluded that the technology was simply too new for anyone to predict when any such aerospaceplane could ever be built, and funding was wound down in 1964. Retrieved from "http://en.wikipedia.org/wiki/Aerospaceplane"

3 Early 1980’s Round One RASV, AMSC, TAV • Manned • Horizontal Takeoff and Landing, HTHL • Single stage, or a near as possible • Fully Reusable, or a s near as possible • SSME / Existing propulsion

• No SSTO designs submitted (or possible) • Propulsion & Structural Fractions too high

Specified preferred solutions, TSTO VTHL solutions ignored

Reusable Aerodynamic Space Vehicle (RASV) The Boeing RASV comprised a ground-based sled to accelerate the aircraft to takeoff speed on a conventional runway, and a delta-winged, piloted orbital vehicle. The RASV was designed to be constructed of conventional refractory metals such as titanium and Rene-41, with the cryogenic liquid hydrogen and liquid oxygen propellants contained within the "hot structure" wing acting as a heat sink to cool the airframe and reduce weight. Powered by two modified SSMEs, the RASV attracted considerable attention from the Air Force, which invested $3 million in the project for technology development in the early 1980s. Advanced Military Space Flight Capability (AMSC) Initiated by AFSC with one-year study contracts awarded to General Dynamics and Rockwell in 1981. Technology studies of small manned spacecraft based on two generic launch concepts: subsonic air launch and "staged" ground launch. Replaced by the Advanced Military Space Technology (AMST) program with a contract awarded to Boeing in January 1984 to determine key aerodynamic and performance parameters associated with air-launching a so-called AMST/TAV orbiter from a carrier aircraft. Transatmospheric Vehicle (TAV) Program begun in mid-1982. Stanley Tremaine coined the term “Transatmospheric Vehicle”, as the craft should be able to operate with equal efficiency both within the atmosphere and in space and be capable of transitioning from space into the atmosphere and back. Phase I of the TAV study began in May 1983 with Battelle Laboratories working with Boeing, General Dynamics, Lockheed, and Rockwell. McDonnell-Douglas submitted its own unsolicited TAV proposal. Phase I ended in December 1983 and resulted in 14 vehicle concepts. Phase II started in August 1984 with a twelve-month contract to Science Applications. In Phase II selected industry concepts were evaluated against alternative solutions such as advanced aircraft and the necessary technologies were further examined with emphasis on determining the military effectiveness of a TAV. The TAV was expected to be the size of a small airliner with a gross liftoff weight of 1 to 1.5 million lbs and using up-rated SSMEs for propulsion in the first generation. A TAV Project office was established in December 1984 under the direction of Lt Col Vince Rausch. By early 1986 the TAV program had been replaced by NASP with the entire TAV staff transferring into the NASP JPO. Science Dawn A classified program begun in 1982 to determine the technical feasibility of a military aerospace plane. Requirements were for a sled-launched horizontal-takeoff / horizontal-landing single stage to orbit (SSTO) launch vehicle powered by a modified SSME with a two-position nozzle. Dry mass was to be 100,000-150,000 lb, takeoff mass 1,2-1,5 million lb with a 10,000 lb payload to a polar orbit from Grand Forks AFB. The craft was to have a turn-around time of 12 hours, be ready to launch within two hours of an alert and have 24 hours orbital capacity. AMSC concepts from Boeing, Lockheed and McDonnell Douglas were hand-picked for further development with Rocketdyne and Air Products as propulsion contractors. The Boeing concept was the RASV vehicle. The McDonnell Douglas proposal was named the Global Range Mach 29 Aerospace Plane, or GRM-29A. It had a down-pointing SSME in the nose to cater for the runway requirements. The Lockheed Zero Length Launch TransAtmospheric Vehicle (ZEL-TAV) used a ramped takeoff with two solid boosters. By 1984 it had become clear that horizontal takeoff was inappropriate use of rocket power, and the program was superseded by Science Realm.

4 Round One Early 1980s Trans Atmospheric Vehicles

Conclusions: Short Comings: No SSTO designs submitted (or possible) Overly constrained Propulsion & Structural Fractions too high Inadequate models with limited physics Level 0-1 analysis applied to systems with very high sensitivities

5 Aerospaceplane deux Round Two 1983-1993 SSTO HTHL, Copper Canyon, NASP

• Manned • Horizontal Takeoff and Landing • Single Stage to Orbit • Fully Reusable

Mandated Solution, Too much money, too little systems engineering

Have Region In 1986 Science Realm was followed by the Have Region program, to complement the ongoing air-breathing work in the NASP program. Main goal was to further develop structures and TPS to reduce risk. Under the program three prototype lightweight structures in scales from 40 to 100% were fabricated from exotic metals, primarily titanium and high-temperature superalloys, to evaluate near-term flight readiness. The cross- sectional structures with integral cryogenic tanks were tested in simulated ascent and re-entry conditions. In tests the Boeing concept was validated and the built but untested Lockheed and McDonnell Douglas designs were classified as partial successes. The test articles were within 3% of required SSTO design weights. Regardless in 1988 it was concluded that the materials developed for NASP were more promising. Total cost of Have Region was around $40 million. Copper Canyon and NASP In June 1983 DARPA initiated the classified Copper Canyon program to investigate the potential military applications of air-breathing hypersonic and single stage to orbit vehicles and technologies with Vince Rausch as project director. Tony DuPont’s initial design from 1983 originated from a NASA study into engine cycles and was a 50,000 lb “F-15 sized“ aircraft. Funded with $6 million for 1983, with Battelle Laboratories doing the main work and initial contracts for airframe work to Boeing, Lockheed and General Dynamics, and propulsion work to Marquardt and GASL. In mid- 1985 the TAV group at the USAF Aeronautical Systems Center became aware of the work being done under Copper Canyon. The TAV studies and contractor designs had concentrated on rocket-based single stage to orbit vehicles rather than the air-breathing vehicles envisioned by Copper Canyon. Soon both groups were discussing a multibillion dollar effort to produce a single stage to orbit aircraft in collaboration with NASA and other DOD agencies to reduce the cost of access to space. In October 1985 the USAF ASC launched the Advanced Aerospace Vehicle (AAV) program to develop advanced hypersonic military aircraft. In December 1985 the National Aerospace Plane (NASP) program was born, a civilian led national effort to develop a single stage to orbit vehicle. NASP JPO was established in January 1986, and president Reagan announced the NASP in his first state of the union address in February 1986. Copper Canyon constituted Phase I of NASP program. The objective of NASP was to develop two flight vehicles with air breathing propulsion from takeoff to orbit. Rocket engines were to be used for final orbital insertion and orbital maneuvers. Airframe, engine module and test facility RFPs were sent out in November 1985. In April 1986 the first design contracts for the NASP program were awarded. Contractors included Boeing, Lockheed, McDonnell Douglas, General Dynamics and Rockwell, with Pratt & Whitney and General Electric for propulsion research. Rocketdyne later contributed to NASP under its own funding. In October 1987, following Phase 2A evaluation, Lockheed and Boeing were dropped from the NASP program, alongside General Electric. Throughout the NASP program the Department of Defense had an 80% share of all money spent on the program. In 1989, during his first week of office as the Secretary of Defense, Richard Cheney terminated the DoD NASP effort. A program review by the National Space Council recommended extending Phase 2 to 1993. In 1991, the National Team program approach combined the resources of the five contractors in a joint-venture partnership to develop a single X-30 concept. Upon completion of the Phase 2D technology development portion in 1993, the technological maturity was deemed not to be at the level required to justify a $15 billion investment to develop two X-30 aircraft in Phase 3, and the NASP was finally cancelled in May 1993. Sources: The Hypersonic Revolution. Volume II. From Scramjet to the National Aero-Space Plane (1964-1986). Richard P. Hallion (Ed.), USAF Aeronautical Systems Division, 1987. The Hypersonic Revolution. Volume III. The Quest for the Orbital Jet: The National Aero-Space Plane Program (1983-1995). Larry Schweikart, Air Force History and Museums Program, 1998. Single Stage to Orbit: Politics, Space Technology, and the Quest for Reusable Rocketry. Andrew J. Butrica, Johns Hopkins University Press, 2003. A Near Term Reusable Launch Vehicle Strategy (http://chapters.nss.org/ny/LongIsland/articles/ANSER.PDF) http://fas.org/irp/mystery/index.html http://www.netwrx1.com/skunk-works/v04.n036 http://www.secretprojects.co.uk/forum/index.php/topic,315.0.html

6 Lets give the rocket 1990-2001 boys a chance Round Three SSTO VTO, DC-X, Venture Star

More mandated solutions, Too much money, too little systems engineering

Venture Star Proposed design for a single-stage-to-orbit reusable launch system by Lockheed Martin. The program's primary goal as a United States federally funded program was to develop a reusable unmanned spaceplane for launching satellites into orbit at a fraction of the cost of other systems that would completely replace theSpace Shuttle. While the requirement was for an unmanned launcher, it was expected to optionally carry passengers as cargo. Venture Star was to be a single-stage-to-orbit vehicle that would take off vertically and land like an airplane. Venture Star was to be a commercial endeavor, and flights would have been leased to NASA as needed. After failures with the X-33 test vehicle, funding was cancelled in 2001.

Science Realm The program was initiated in 1984 as a follow-on to Science Dawn. In contrast to Science Dawn, which stressed horizontal takeoff, Science Realm investigated vertical takeoff SSTO designs, capitalizing on the high thrust-to-weight ratio of a rocket engine. During the program structural test articles were designed based on Science Dawn designs. The cost of Science Dawn and Science Realm together was about $20 million.

The DC-X, short for Delta Clipper or Delta Clipper Experimental, was an unmanned

7 prototype of a reusable single stage to orbit launch vehicle built by McDonnell Douglas in conjunction with the United States Department of Defense's Strategic Defense Initiative Organization (SDIO) from 1991 to 1993. After that period it was given to NASA, which upgraded the design for improved performance to create the DC-XA. Contents [hide] 1 Background 2 Design 3 Flight testing 4 The future of the DC-X 5 See also 6 References 7 External links [edit] Background According to writer Jerry Pournelle: "DC-X was conceived in my living room and sold to National Space Council Chairman Dan Quayle by General Graham, Max Hunter and me." According to Max Hunter, however, he had tried hard to convince Lockheed- Martin of the concept's value for several years before he retired.[1] Hunter had written a paper in 1985 entitled "The Opportunity", detailing the concept of a Single- Stage-To-Orbit spacecraft built with low-cost "off-the-shelf" commercial parts and currently-available technology[2], but Lockheed-Martin was not interested enough to fund such a program themselves. On February 15, 1989, Pournelle, Graham and Hunter were able to procure a meeting with Vice-President Dan Quayle. They "sold" the idea to SDIO by noting that any space-based weapons system would need to be serviced by a spacecraft that was far more reliable than the Space Shuttle, and offer lower launch costs and have much better turnaround times. Given the uncertainties of the design, the basic plan was to produce a deliberately simple test vehicle and to "fly a little, break a little" in order to gain experience with fully reusable quick-turnaround spacecraft. As experience was gained with the vehicle, a larger prototype would be built for sub-orbital and orbital tests. Finally a commercially acceptable vehicle would be developed from these prototypes. In keeping with general aircraft terminology, they proposed the small prototype should be called the DC-X, X for "experimental". This would be followed by the "DC-Y", Y referring to pre-run prototypes of otherwise service-ready aircraft. Finally the production version would be known as the "DC-1". The name "Delta Clipper" was chosen deliberately to result in the "DC" acronym, an homage to the famous DC-3 aircraft, which many credit for making passenger air travel affordable.[citation needed] [edit] Design The DC-X was never designed to achieve orbital altitudes or velocity, but instead to demonstrate the concept of vertical take off and landing. The vertical take off and

7 landing concept was popular in science fiction films from the 1950s (Rocketship X-M, Destination Moon, and others), but not seen in real world designs. It would take off vertically like standard rockets, but also land vertically with the nose up. This design used attitude control thrusters and retro rockets to control the descent, allowing the craft to begin reentry nose-first, but then roll around and touch down on landing struts at its base. The craft could be refueled where it landed, and take off again from exactly the same position — a trait that allowed unprecedented turnaround times. In theory a base-first re-entry profile would be easier to arrange. The base of the craft would already need some level of heat protection to survive the engine exhaust, so adding more protection would be easy enough. More importantly, the base of the craft is much larger than the nose area, leading to lower peak temperatures as the heat load is spread out over a larger area. Finally, this profile would not require the spacecraft to "flip around" for landing. The military role made this infeasible, however. One desired safety requirement for any spacecraft is the ability to "abort once around", that is, to return for a landing after a single orbit. Since a typical low earth orbit takes about 90 to 120 minutes, the Earth will rotate to the east about 20 to 30 degrees in that time; or for a launch from the southern United States, about 1,500 miles (2,400 km). If the spacecraft is launched to the east this does not present a problem, but for the polar orbits required of military spacecraft, when the orbit is complete the spacecraft overflies a point far to the west of the launch site. In order to land back at the launch site, the craft needs to have considerable cross-range maneuverability, something that is difficult to arrange with a large smooth surface. The Delta Clipper design thus used a nose-first re-entry with flat sides on the fuselage and large control flaps to provide the needed cross range capability. Experiments with the control of such a re-entry profile had never been tried, and were a major focus of the project.[citation needed] Another focus of the DC-X project was minimized maintenance and ground support. To this end, the craft was highly automated and required only three people to man its control center (two for flight operations and one for ground support). In some ways the DC-X project was less about technology research than operations. [edit] Flight testing The Delta Clipper Advanced Construction of the DC-X started in 1991 at McDonnell Douglas' Huntington Beach facility. The aeroshell was custom-constructed by Scaled Composites, but the majority of the spacecraft was built from "off the shelf" parts, including the engines and flight control systems. The DC-X first flew, for 59 seconds, on 18 August 1993. It flew two more flights 11 September and 30 September, when funding ran out as a side effect of the winding down of the SDIO program. Apollo astronaut Pete Conrad was at the ground-based controls for some flights.[3] Further funding was forthcoming, however, and the test program re-started on 20 June 1994 with a 136 second flight. The next flight, 27 June 1994, suffered an inflight

7 (minor) explosion, but the craft successfully executed an abort and autoland. Testing re-started after this damage was fixed, and three more flights were carried out on 16 May 1995, 12 June, and 7 July. On the last flight a hard landing cracked the aeroshell. By this point funding for the program had already been cut, as a side effect of the winding down of the SDIO program, and there were no funds for the needed repairs. NASA agreed to take on the program at this point. In contrast to the original concept of the DC-X demonstrator, NASA applied a series of major upgrades to test new technologies. In particular, the oxygen tank was replaced by a lightweight (alloy 1460 equivalent of alloy 2219) Al-Li tank from Russia, and the fuel tank by a newer composite design. According to Bob Hartunian (former McDonnell Douglas and Boeing cryo-tank specialist), the Russian-made tank was poor quality, had "16- inch/40.6-cm long weld defects, and there were other issues that, according to U.S. standards, would prevent it from flying." [4] The control system was likewise improved. The upgraded vehicle was called the DC- XA, renamed the Clipper Advanced/Clipper Graham, and resumed flight in 1996. The first flight on 18 May 1996 resulted in a minor fire when the deliberate "slow landing" resulted in overheating of the aeroshell. The damage was quickly repaired and the vehicle flew two more times on 7 and 8 June, a 26-hour turnaround. On the second of these flights the vehicle set its altitude and duration records, 3,140 meters and 142 seconds of flight time. Its next flight, on 7 July, proved to be its last. During testing, one of the LOX tanks had been cracked. When a landing strut failed to extend due to a disconnected hydraulic line, the DC-XA fell over and the tank leaked. Normally the structural damage from such a fall would constitute only a setback, but the LOX from the leaking tank fed a fire which severely burned the DC-XA, causing such extensive damage that repairs were impractical.[5] In a post-accident report, NASA's Brand Commission blamed the accident on a burnt- out field crew who had been operating under on-again/off-again funding and constant threats of outright cancellation. The crew, many of them originally from the SDIO program, were also highly critical of NASA's "chilling" effect on the program, and the masses of paperwork NASA demanded as part of the testing regimen.[citation needed] NASA had taken on the project grudgingly after having been "shamed" by its very public success under the direction of the SDIO. Its continued success was cause for considerable political in-fighting within NASA due to it competing with their "home grown" Lockheed Martin X-33/VentureStar project. Pete Conrad priced a new DC-X at 50 million dollars, but NASA decided not to rebuild the craft in light of the budget constraints.[5] Rather, NASA focused development on the Lockheed Martin VentureStar which it felt answered some criticisms of the DC-X; specifically the requirement that many NASA engineers preferred the airplane-like landing of the VentureStar over the vertical landing of the DC-X. First flight First landing. The yellow exhaust is due to the low throttle settings, which burns at

7 lower temperatures and is generally "dirty" as a result.[citation needed] Height 12 m (39' 4")[citation needed] Diameter 4.1 m (13' 5")[citation needed] Dry mass: 9,100 kg[citation needed] GLOW: 18,900 kg Propellants: Liquid oxygen and liquid hydrogen Engines: Four RL-10A-5 rocket engines Engine thrust: 6,100 kgf[citation needed] Reaction controls: four 440 lbf (2,000 N) thrust gaseous oxygen, gaseous hydrogen thrusters

7 1990s Developing Engineering Models

• Design & Analyses • Aerothermal • Thermal Protection • Thermal Management • Scramjets • Inlets & nozzles • Weights • Rockets • Subsystems • Operations Models

• Cost LRC

Finally some systems engineering models

8 Maintenance Modeling 2002-3

HRSI Tile Removals 0.25 Mods Ground 0.2 Access TPS Component Man-hours 0.15 Total . AFRSI Blankets Attach & Gap Simulation & Thermal Barriers Probability 0.1 Rollup model Flight Inspections .

0.05 Gap Fillers Total TPS Man-hours HRSI Tile 0 0 10 20 30 40 50 60 70 Tile Removals HoursTriangular to Probability Remove Distribution & Replace Tile 4 4 4 4 4 4 0 5000 1 10 1.5 10 2 10 2.5 10 3 10 3.5 10 Total TPS Manhours .

Man-hours calculated for each maintenance action for each 120 140 160 180 200 Man Hours / Tile component

Maintenance Model is sensitive to Design and Technology 9

9 2003-4 Operations Models

• Problem: Shuttle Turn-Time (3-4 months) and RLV Turn-Time Goal (8-48 hours)

• Solution: Link conceptual designs to Turn-Time and maintenance actions

Shuttle Ground Ops Aircraft Ground Ops Innovative Design, Advanced Technologies

& Thorough Development is needed 10

10 Cost Modeling Cost Analysis: No ELV Development and No HLV Engine Development

HLV: Low Risk and Low Cost Path to RLVs 11

A real world scenario was examined that included no development costs for ELVs. Since the vehicles are currently in use today, only the DOC was important. No 2nd stage engine development was done for the hybrid system since the engine is expendable and current exist. The RLV undergoes complete development and production. Each vehicle was sized for a payload mass of 15,000 lbs and the system life was determined to be 20 years. Using this real world scenario, it is apparent that the expendable system remains preferable for 190 launches over a 20 year system life. That roughly corresponds to nine launches a year over 20 years. The hybrid system is preferred for launches of 190 to 270 launches and the reusable system preferred for launches greater than 270 over 20 years. The U.S. military launched a total of six missions in 2005. However, it is understandable that launch rates are going to increase. The hybrid system would offer lower costs for launch rates of 10 or more a year and the reusable for launch rates great than 15.

11 Ops Driven Design Summary 2004

• Simplified system design • Launch systems are as important as flight systems. • Easy access, removal, replacement • Elimination of toxic fluids • Rugged Thermal Protection System (TPS) • Minimize TPS area • Minimize number of fluids ( HC, O2, N2 ) • Eliminate or minimize high pressure hydraulics & pneumatics

12 Round Four 2003 Comparative Studies

Rocket Systems (R) Air Breather Systems (AB) Vertical Takeoff (VT) Horizontal Takeoff (HT)

Vertical Landing (VL)

Finally some systems engineering comparative studies.

13 Reusable Launch System (RLS) Missions of Interest

(ASC/FB) Rapid Global Strike ~2020+ IOC SOS FSS ACC (Sub-Orbital Strike) (Future Strike System) $20B+ Dev Acq

Routine, Flexible Mil. Space Trans. (SMC/XR) ~2015+ IOC RLS SOV MSP AFSPC (Space Op. Veh.) (Mil Space Plane) AFSPC SMC ASC Low Cost Space Transportation (NASA MSFC) AFRL ~2015+ IOC NASA & (SpaceSTS Trans Sys) 2 3 Gen NASA $20B+ Dev Acq

Multiple Missions are needed to justify

9/24/2004Reusable Launch Systems 14

• The Reusable Military Launch System (RMLS) team is pursuing systems which could meet all, or most, of the above needs. •Initial Operating Capability (IOC) and Technology Availability Dates (TAD) indicate direction given to associated mission needs. •Finding enough missions to justify a RLV is a major problem

14 Figures Of Merit

GOOD FIGURES OF MERIT: – Empty Weight • Indicates development and acquisition costs – Complexity • Indicates development and acquisition costs and risks – Wetted Area • Indicates TPS inspection and maintenance (ops) costs – Uncertainties & Sensitivities • Indicates technical and design readiness – Growth Factor • Directly measures closeness to the technical limit – Maintenance Man-hours per launch • Main contributor to operating and life cycle cost • Launch Weight is NOT a good Figure Of Merit (FOM) – It does not differentiate between propellants and hardware

9/24/2004 15

Ultimate figures of merit such as robustness, flexibility, risk, safety, mission cost and total life cycle cost are not very useful during the wide ranging conceptual phase of systems design and analysis. Needed are a few figures of merit which are good indicators for the above. This paper will make use of the following: Empty Weight Empty weight is a good indicator for development and acquisition costs. It is widely used at this stage of analysis, and is often used along with material types and complexity. Complexity Overall system complexity impacts development and acquisition costs and risks. Measuring complexity is not very quantitative and can be subjective, but is worth noting. One typically counts the number and type of major subsystems and the complexity of their interactions. Maintenance man-hours is a reasonable measure of this as well, though it does bring in the quality of the development process, in that the reliability of the subsystems are part of the equation. Wetted Area Wetted area is a good indicator for Thermal Protection System (TPS) related costs. TPS is a major contributor to maintenance man hours and turn time. Uncertainty Uncertainty estimates in conjunction with growth factors indicate technology readiness level and help set appropriate management margins for system development. The impacts on various uncertainties within the system are estimated using a Monte Carlo technique. Growth Factor There are many different growth factors. The one we are interested in is the empty weight growth factor. It is defined as the growth in system empty weight needed to restore full system flight performance in response to a change in weight. It is obtained by differentiating the system sizing equation with respect to a change in weight. It grows asymptotically as the system approaches its performance limit. High growth factors combined with uncertainties can yield extreme variations in the final size. This is especially true of HTHL SSTO systems. Maintenance Man-Hours per launch cycle Average total man-hours of maintenance needed to prepare the complete system for its next launch is a major contributor to operating and life cycle costs. Modeling maintenance man-hours has been a major effort of team member Brendan Rooney and is reported in Reference 1. As a final observation, gross liftoff weight is NOT a particularly good figure of merit for this type of wide ranging comparative analysis, because it does not differentiate between propellants and hardware. Hardware is extremely expensive to develop, acquire and maintain, whereas propellants, even hydrogen, cost practically nothing by comparison.

15 Some Study Results Staging Velocity ~ 8600 fps Ideal Staging DV ~ 12,500 fps Smaller Wing Payload – NOT returned

Staging Velocity ~ 7300 fps

C1-270nm-E C2-100nm-E C3-100nm-E C4-100nm-E C5 100nm-E Wg: 2,345,344 2,345,344 2,345,344 2,345,344 284,024 We: 273,882 274,442 256,070 248,310 49,422 Wpyld: 9,500 18,550 36,970 45,340 1,500 Len: 121 ft 128 ft 128 ft 128 ft 1 block = 20 ft 67 ft

C1 & C2 payload dimensions 15x30 C3 & C4 payload dimensions 12x30 => (23% packing density difference for these cases) 22% decrease in wing-span (C3 -> C4) 39% decrease in trapz-wing-area (C3 -> C4) Relates to decreases in structural aerosurface wt, TPS, LG, and actuation.

Fineness-Ratio difference on C1 is 0.2% of We. Possibly better aerodynamics for larger Fineness-Ratio. (5.4 vs 6.0)

16 Alternative System Concepts

TSTO Rocket Systems: serial burn, stage 7000 fps, fly-back booster LH LH LH/lox engines in both stages RP RP RP1/lox engines in both stages M M Methane engines and subsystems in both stages M M w Methane engines in both stages, and Inconel structures Multi-Stage Hybrid Rocket Systems: smaller reusable booster RP RP High performance RP1/lox engines used on upper stage RP LH LH/lox upper stage RP S2 Two solid rocket upper stages

VTHL SSTO Rocket Boosted Scramjet Systems: xR-LHSJ-LHR V HC 2D RP1/lox boost rockets, 2D scramjets. V HC Inw RP1/lox boost rockets, One Inward turning scramjet V LH Inw LH/lox boost rockets, One Inward turning scramjet.

VTHL TSTO Rocket Boosted / Scramjet Systems: Glide-back Booster V HCR /2D RP1/lox booster with a 2D scramjet upper stage V HCR /SJ RP1/lox booster with an inward turning scramjet upper stage

HTHL SSTO Scramjet Systems: R-LHSJ-LHR H LH 2D LH/lox boost rockets with 2D scramjets H LH Inw LH/lox boost rockets with inward turning scramjets H HC Inw RP1/lox boost rockets with inward turning scramjets

HTHL TSTO Scramjet / Rocket Systems: xx-LHSJ /10k/ LHR TBCC /R HC turbojet to Mach 3.5+ and 2D LH scramjets on the 1st stage RBCC /R LH/lox rockets to Mach 2.5+ and 2D LH scramjets on the 1st stage We examined all of these with a consistent methodology 9/24/2004 17

Numerous rocket, scramjet and turbine propulsion combinations are considered. TSTO and SSTO systems are looked at, as are vertical and horizontal takeoff systems. This paper reports on 17 different launch system concepts, and more are being added. All systems are fully reusable except the hybrid systems. A brief description of each of the systems follows. More details can be found in the references. The system’s identifying nomenclature is used in the included graphs.

17 Gross Weight (klbs)

1859

1531 1341

1604 1255 1310 1148 1473 972 961 1078 1157 923 1360 823 907 1216 943843 694 725 1263 1034 1111 903 703 709 836721 585 1180 793 903 602 572 1019 750630 503 649 719 822 531 465 944 539 442 585 667 648 388 469 860 525 587 479 3 430 3 c_ 3 a 2 _ D 2 a 2 b _ D 2 _ IN b 2 S6 _ D 0% 8 IN +1 b 7c _ 7 2 SSTOH e a D Bas 6 b 6b_IN _ SSTOV THL 2 0% 4c D -1 4 TSTOH THL a Growth % TSTOV THL 9/24/2004TSTO THL 18 Rocket

18 Empty Weight (klbs)

422 426

417 411 351

365 359 349 320 291 319 320 302 238 257 199 218 284 280 255 165 177 247 243 151 168 235 189 220 137 140 123 214 189 221 164 104 114 113 174 142 195 147 89 98 158 126 177 129 76 146 110 160 114 135 99 3 3 c a _ 87 3a_IN _ 2 2b_2D 2 D 2b_IN D 0% 8 +1 b 7 _ c 2 e 7a D SSTOH Bas 6 6 b SSTOV b _ THL 4c _ 2 10% I D - 4a N TSTOH THL Growth % TSTOV THL 9/24/2004TSTO THL 19 Rocket

19 Wetted Area (kft2)

64.8

43.9 40.3 39.8

33.8 30.8 32.4 34.8 35.2 27.9 23.2 25.3 28.1 33.2 32.4 25.5 21.0 20.8 31.3 30.4 21.8 23.5 22.9 18.5 18.6 28.5 19.1 28.9 23.1 29.1 21.3 16.7 16.8 21.1 26.4 27.4 21.2 19.5 27.6 19.4 15.5 15.6 19.9 26.0 20.3 18.2 18.1 14.2 24.7 19.1 16.9 18.6 18.4 16.1 17.4 23.3 3 17.7 15.0 3 c_ 3 a 2 a _ D 2b_2D _ 2 2 I D b N _ 0% 8 IN +1 b 7 _ c 2 e 7a D SSTOH Bas 6 b 6b_IN _ SSTOV THL 2 0% 4c D -1 4a TSTOH THL Growth % TSTOV THL 9/24/2004TSTO THL 20 Rocket

20 Boost-back vs. Fly-back 2005 Dev & Prod Cost Estimation

25

Flyback Development Boostback Development Flyback Production 20 Boostback Production

21.7%

15

$40M Upper Stage Minimum DPM Cost Points 10 Cost ($B FY06)

120,000 5

100,000 15% Dry Weight 0 Margin 3456789 Other Subsystems 80,000 15% Dry Weight Staging Mach Number Margin TPS Other Subsystems Air Breathing 150 60,000 Hardware Rocket Engine and Propellant Feed Rocket Engine and 100 ih lm) (lbm eight W 40,000 Propellant Feed Ascent Trajectories Boostback Landing Gear Landing Gear Other Fuselage Other Fuselage 20,000 Structure Structure 50 Flyback Propellant Tanks Propellant Tanks Altitude (nmi) Aero Surface Aero Surface 0 0 Boostback Flyback 0 50 100 150 200 250 300 350 400 Downrange (nmi) 21

21 Advanced Launch Architecture

Multi mission capable: - Prompt Global Reconnaissance and Strike - Prompt access to space - Prompt rendezvous to orbits - Lower Costs as missions expand

22 Launch System Elements

• System Components: Reusable Upper stage – Reusable Booster • Needed for all options Scramjet-rocket – Expendable Upper stage Upper stage • Prompt Global Recce-Strike • Prompt access to space Reusable – Reusable Rocket Upper stage Hybrid Booster Launch • Lower costs with more missions Config. – Reusable Scramjet Upper stage • Prompt Rendezvous to orbit

The Reusable Booster is the critical First step

23 Some Conclusions

• Numerous reusable and partially reusable rocket systems are attractive and technically achievable now. Viability depends on launch rates.

• Technology investment, design and system development should be focused on operability.

• Horizontal launch increases the technical risk, turn-time and cost of launch systems.

• “Aircraft-like” operations for “access to space” does not necessarily mean turbine based or horizontal takeoff systems.

• Airbreathing propulsion for “access to space” should focus on vertically launched rocket/scramjet systems with an eye to SSTO.

9/24/2004 24

Numerous reusable and partially reusable rocket systems are attractive and technically achievable now, but sufficient numbers of annual missions and reduced costs are needed to justify their development. They are all two or more staged systems, and are vertically launched. The partially reusable systems are comprised of a reusable booster with expendable upper stages. Vehicle turn times of one to two weeks appear to be achievable with existing technology by using “ops-focused” design and development practices. The desired time of 24-48 hours (to be more “aircraft like”) is beyond the state of the art but is within reach with a reasonable level of, again, “ops-focused” technology investment. Further reductions of turn time and operations costs will require technical advances to support development of durable, operable thermal protection systems, rocket engines, and fluid related subsystems. Maintenance and turn time is dominated by these subsystems. Technology development, design and system development must be focused on operability. A thorough (more aircraft like) development program will be necessary to obtain the levels of operability desired. Such a program could easily extend up to 10 years. Systems with considerable design margin make development easier. To this end, it is paramount that we keep our requirements lean, focus on simple systems and optimize them for high operability. Horizontal launch does not improve the turn time of launch systems, it increases it. This is due to the increased maintenance associated with their larger size. Times to mate, transport and fuel should be similar for both horizontal and vertical launch, while vertical launch pad mating, erection and fueling can be made small relative to the maintenance times of large horizontally launched systems. Systems such as Zenit have shown launch times of a few hours by automating the procedures. “Aircraft-like” operations for “access to space” should not necessarily imply turbine based horizontal takeoff systems. Conventional wisdom is sometimes wrong. These launch systems fare very poorly against vertically launched rocket and combined rocket/scramjet based systems. Horizontally launched TBCC/rocket systems are among the worst, being larger, having heavier empty weights and much greater wetted areas, including extensive amounts of actively cooled area. They are more complex, contain three different types of engines, and will have significantly larger amounts of maintenance. Having said all of this, there are hypersonic cruise missions where TBCC may be the engine of choice, but not for pure access to space. Airbreathing propulsion for “access to space” should focus on vertically launched rocket/scramjet systems with an eye to SSTO. Good scramjet performance in the Mach 10-15 régime, light weight integral tanks and advanced thermal protection systems are critical in achieving this goal. It is very difficult for a TSTO airbreathing concept to compete with a TSTO rocket system; having said that, a TSTO rocket boosted scramjet would be a lower risk first step towards SSTO. This system could launch considerable payload using the small hybrid class booster, and pave the way for SSTO. Above all we need to keep these scramjet designs as simple as possible without giving up too much margin.

24 What I learned about Systems Engineering

• Do Analysis of Alternatives and heed the results. • Unified System Models are required – Proper Models are critical !!! – If you don’t have good models, don’t commit large amounts of money. • Consistency in analysis of alternatives is more important than absolute accuracy. (apples to apples) – individual designs done by separate organizations will raise more questions than they answer. (apples & oranges) – Alternative comparisons done by separate organizations will be much more useful.

If you’re not consistent building models and running trades, you don’t know much

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