A Near Term Reusable Launch Vehicle Strategy

Transcription

1 A Near Term Reusable Launch Vehicle Strategy Ramon L. Chase Warren Greczyn Leon McKinney February 2003 (update) 2900 South Quincy Street Arlington, VA

2 Introduction Provide data that could be used in the formulation of 2008 operational prototype reusable military space plane candidates Consistent with Air Force aircraft operational concepts Mature into a full operational capability by 2010 Space launch mission capabilities Military space plane and space launch legacy ( s) What was accomplished in previous programs How might it be applied today to a near term reusable launch vehicle 2

3 Program Legacy Air Force Aerospace plane and DynaSoar programs CIA Isinglass program NASA and Air Force4 single stage to orbit studies Air Force Military Crews in Space study AF HQ (RDSL) mission analysis and cost benefit studies Air Force Advanced Manned Space-Flight Capability study Air Force Transatmospheric Vehicle study Air HQ (RDSL) Science Dawn program Air Force Science Realm program DARPA Copper Canyon study Air Force Have Region program DOD-NASA National AeroSpace Plane (NASP) program SDIO DC-X program NASA & Lockheed-Martin X-33 program NASA Space Launch Initiative program 3

4 Military Space Plane Legacy of the s AMSC Isinglass Isinglass: rocket-powered air launched reconnaissance aircraft AMSC: AF Advanced Manned Space flight Capability program Science Dawn: MSP program after down select from AMSC concepts Science Realm: design of structural test articles from SD designs Have Region: manufacture and test of SR subscale structural test articles Copper Canyon (DARPA): air-breathing SSTO concept NASP: air-breathing SSTO Concept Science Dawn Science Realm Copper Canyon Have Region NASP 4

5 Isinglass Isinglass: Mach 17, rocket powered, air-launched reconnaissance aircraft. RS-129: Staged combustion hydrogen-oxygen rocket component development program Light weight structures: Diffusion bonded titanium honeycomb 5

6 Reusable Launch Vehicle Legacy of the 1990s: What Happened? Have Region NASP 1988, Change of Administration, new priorities 1988, Have Region completed, no follow-on 1992, NASP X-30 program cancelled 1996, NASA X-33 rocket powered SSTO technology demonstration program initiated 2001, X-33 program cancelled 2001, NASA Space Launch Initiative Program 2001, NASA-Air Force near term activity X-33 NASA SLI 6

7 Reusable Launch Vehicle Technology Legacy Summary Isinglass Light weight structure and TPS concepts applicable to reusable launch vehicle concepts Science Dawn 3 Military space plane concepts In-depth preliminary designs by hand picked design teams Upgraded SSME performance, 109 and 115 %, horizontal operation, 2-position nozzle Have Region 3 structural concepts, 2 partial successes, 1 validated concept Stainless steel hydrogen tank liner Propellant utilization fraction of 0.88 NASP X-33 New materials Composite hydrogen tank, propellant utilization fraction of 0.74 Carbon-carbon aero surfaces High temperature heat pipe sharp leading edges Linear aerospike engine development 7

8 Attributes of a Near-Term Cost Optimized Reusable Launch Vehicle Design Concept Design Reference Mission Minimize ideal velocity requirements Design Concept Simple Configuration, validated structural concept, maximum use of existing subsystems, components and facilities Propulsion Existing engine, minimize modification requirements, long life components, boost pumps and low pressure propellant tanks, common APU, RCS and main propulsion propellants Structural Design Current materials, low cost manufacturing processes, flat surfaces, hot structure (no TPS, no insulation, no tiles), integrated load bearing structure and tankage, flat bulkheads, low propellant pressurization, structural margin Operational Concept and Infrastructure Airplane like operations (life cycle maintenance), built in test and status monitoring, automated fueling, containerized payloads, co-location at existing operating bases, horizontal take-off, self ferrying 8

9 Low Cost, Near Term RLV Design Strategy Design reference mission: First stage performance requirement, global range boost-glide-skip 3,000 fps less than orbital velocity First stage usable propellant fraction approximately 0.82 No first stage orbital maneuvering propulsion system req d Vehicle design characteristics: Two stage, 1, 200,000 GTOW Operational prototype in 2008, full operational capability in ,000 lb in low earth Polar orbit. Full abort capability with two engines, or linear aerospike engine configurations Propulsion 2 modified SSMEs with 2-position nozzle, linear aerospike, or new Pratt & Whitney engine Self ferry capability with removable jet engine pods 9

10 Low Cost, Near Term RLV Design Strategy Continued Structures and materials Validated structural concept at 0.88 propellant utilization fraction Hot structures, integral propellant tanks(inherent in design), flat bulkheads Low pressure propellant tanks plus boost pumps Wet fuselage and wing Isinglass diffusion bonded honeycomb Current materials NASA Ames fight tested sharp leading edge materials Operation characteristics: Single operating base Horizontal maintenance, launch preparation and launch (no launch pad complexes ) Automated propellant loading Containerized payloads 10

11 Where Do We Stand Today? Design reference mission options: Low Earth orbit (current focus) Geo transfer orbit Global range boost-glide-skip Vehicle Design: Shift in emphasis from SSTO to TSTO concepts Potential shift from performance optimized to minimum cost designs Propulsion: Modified Shuttle main engine and the RL-10 Other existing engines (including air-breathing), or near term options Foreign engines, Russian and Japanese rocket engines Partially developed linear aerospike engine New more reliable, long life engine concepts Structures and Materials McDonnell Isinglass structural concept validated Boeing Have Region structural concept ( hot structure) validated NASP composite hydrogen tank, sharp leading edges, C-C aero surfaces high temperature materials and composites 11

12 What Might Be Possible by 2008 Change in paradigm from launch vehicle to aircraft operations Vehicle design options Vertical take-off two stage to low Earth orbit, multiple bases A subsonic horizontal take-off air-breathing two-stage to low Earth orbit, multiple bases A horizontal take-off assisted single stage global reach with a two stage to low Earth orbit plus a geo transfer orbit option, existing Air Force bases Propulsion Modify existing engines to meet vehicle design requirements Structures Incorporate McDonnell diffusion bonded honey comb into Boeing hot structures concept. Update materials and manufacturing methods Use NASP aero surfaces Use NASA AMES sharp leading edges Update materials in B-58 type landing gear design Subsystems Modified existing subsystems and avionics 12

13 Paradigm Evolution 1950s 1960s 1970s 2000s 2025 Thor, Atlas, Titan Expendable Vertical Launch Build-Up on Pad Saturn Expendable Vertical Launch Build-Up Off Pad and Roll Out Space Shuttle Partially Reusable Vertical Launch Build-Up Off Pad and Roll Out Advanced Launch Vehicles Fully Reusable Horizontal Launch Hangar Service/ Taxi Out 13

14 A Paradigm Shift Now Current Paradigm Continues Two-Stage-to-Orbit Vertical Take-Off Launch Vehicle-Based Ground Operations Rocket Powered Alternative Paradigm Begins Now Two-Stage-to-Orbit Horizontal Take-Off Military Airplane-Based Operations Rocket Powered X Vision Single-Stage-to-Orbit Horizontal Take-Off Fully Reusable Airplane-Based Operations Combined Cycle Propulsion 14

15 Reusable Launch Vehicle Capabilities Horizontal take-off and landing Adverse weather, day and night operations One mission per day sustainable surge capability Twenty four hour turnaround Twelve hour call up from post-flight readiness Two hour flight readiness from alert Co-location with military Military aircraft ground and flight operations Centralized life cycle maintenance Self ferry capability Twenty four hour on-orbit residence capability 15

16 2008 Vehicle Weight Summary* Structure 85,000 lb Body Wing Flight Compartment Vertical Tail Propulsion 21,000 lb Main Engine(s) Ancillary Systems Propellant Feed System Helium Pressurization System Subsystems 18,500 lb Reaction Control System Landing Gear Prime Power Electrical Distribution and Control Aero Surface Controls Avionics Environmental Control Margin 10,500 lb Residuals, Reserves, In-flight Losses, RCS Propellants 9,000 lb Payload (Grand Forks, Polar orbit) 23,000 lb Inert Weight (Including Payload) 167,000 lb Ascent Propellants (including in-flight reserves ) 1,060,000 lb Gross Take-Off Weight 1,227,000 lb Useable Propellant fraction: *Science Dawn data base, Global Range Boost-Glide-skip flight Profile 16

17 Accelerated Program Schedule Summary* Calendar Year Program Milestones Start Final Configuration System PDR 90% Drawing Release Rollout First Flight Aerospace Plane Assist System (Risk Mitigation, Back-Up) Manufacturing Facilities Structure PDR Fabricate and Assemble Build Scale Test Model Decision to Proceed First Flight Launch Site Construction Complete * Based on Science Dawn proposal data, details available 17

18 Reconstitution and Post Flight Maintenance* Minutes Full Stop on Runway Post Flight Operations (19 Technicians, Specialists and Manager) Tow to Maintenance, Refurbishment and Readiness Facility Post Flight Safing (37 Technicians, Specialists, Foreman and Manager) Payload Operations (20 Technicians, Payload Specialists, Foreman and Manager) Flight Readiness Start Refurbishment Operations (38 Technicians, Specialists, Foreman and Manager) Tow to Storage Facility Tow to Readiness Facility Payload Operations (17 Technicians, Payload Specialists, Foreman and Manager) Readiness Operations (37 Technicians, Specialists, Foreman and Manager) Tow to Fueling Facility Propellant Loading (Automated Fast-Fill Operation) Tow to Runway * Boeing/SAC estimate based on detailed subsystem analysis Engine Start, Brakes Off, Begin Take-Off Roll 18

19 Post Flight Ground Operations Jet Fuel Storage (Underground) Space plane Mission Control N Maintenance, Refurbishment and Readiness Facility Payload Storage Helium Storage Space plane Storage Shelters Space plane lands and comes to stop on runway Space plane is towed to Maintenance, Refurbishment and Readiness Facility Post flight safing operations are performed on space plane Fuel purge/vehicle cool-down Vehicle Inspection Payload Operations Remove Payload Container Strip Alert Shelter ATC Tower Conventional Aircraft Hangars Refurbishment Operations Checkout and replace/ repair Tow to Storage Facility activate status monitoring system Helium Storage Fueling Facility LOX, LH2 Storage (Underground) Strip Alert Shelters LOX, LH2 Production Facility 19

20 Flight Readiness Ground Operations Jet Fuel Storage (Underground) Strip Alert Shelter Spaceplane Mission Control N Maintenance, Refurbishment and Readiness Facility ATC Tower Payload Storage Helium Storage Conventional Aircraft Hangars Spaceplane Storage Shelters Upon receipt of ATO, flight readiness operations are initiated. Space plane is removed from storage and towed to Maintenance, Refurbishment and Readiness Facility. Mission flight data base generation is initiated. Payload inspection, checkout and repair operations conducted. Payload containerized and moved to the flight readiness facility for insertion into the space plane. Space plane inspection, checkout and repair operations conducted. Load payload and place mission data into space plane computers. Close out space plane. Transfer to on-board power. Strip Alert Shelters Helium Storage Fueling Facility LOX, LH2 Storage (Underground) LOX, LH2 Production Facility Tow space plane to automated fueling facility. Conduct chill-down operation and load fuel. Tow vehicle to strip alert facility with automated topping. Disconnect fuel topping and start engines. Taxi to runway and takeoff. 20

21 Near-Term RLV Self-Ferry landing Click on Movie to Play 21

22 Near-Term RLV Returning From a Mission Click on Movie to Play 22

23 Near-Term RLV Fueling Click on Movie to Play 23

24 Near-Term MSP Taking Off Click on Movie to Play 24

25 Accelerated Program Schedule Summary* Calendar Year Program Milestones Start Final Configuration System PDR 90% Drawing Release Rollout First Flight Aerospace Plane Assist System (Risk Mitigation, Back-Up) Manufacturing Facilities Structure PDR Fabricate and Assemble Build Scale Test Model Decision to Proceed First Flight Launch Site Construction Complete * Based on Science Dawn proposal data, details available 25

26 Program Cost Assumptions 3 Operational prototype aerospace planes Single operational base 7-10 launches per year Fuel manufacturing and storage facilities Logistics support (25% of flyaway costs) 5 year operational period ( ) Second stage not included ( Plus $2 B ) New rocket engine not included( Plus $2 B ) 26

27 Model 2008 ROM Budgetary Funding (2002 Dollars in Millions) Development Non-Recurring Airframe and Launch Assist $3,000 Avionics 300 Engines - SSME Ground and Flight Test 1,000 - Total Development Non-Recurring $4,570 Development Hardware/Facilities/Operations Air Vehicle (3 Operational Prototypes) - Air Frame Manufacture 3,000 - Avionics SSME GE CF Total Air Vehicle $3,820 Operational Hardware/Facilities - Fuel Facilities Logistics Support 1,000 - Operational Cost Total Operational Hardware/Facilities $2,060 Program $10,450 27

28 Summary Four key design parameters determine the technical feasibility of a reusable launch vehicle. They are: 1. The vehicle design concept, 2. The reference design mission, 3. Propulsion system performance, and, 4. The structural design concept (determines the usable propellant fraction achievable). Near term design concepts include both assisted SSTO and TSTO options. After careful review three years ago the AF SAB considered SSTO to be to high a risk to be considered as near term. Design reference mission options include: 1. A low Earth orbit, or 2. A Geo transfer orbit. Propulsion choices include: 1. A subsonic jet engine. 2. A supersonic jet engine, 3. A supersonic combined cycle turbo- ramjet, and 3. Several rocket engine options. Two structural design concepts have been validated. They are the Douglas Isinglass and the Boeing Have Region. Two additional Have Region test articles were designed and built, but not tested. They are the Lockheed Martin and McDonnell Douglas (now Boeing). 28

29 Summary Continued A low earth orbit SSTO capable reusable launch vehicle is considered high risk in the near term. Currently available propulsion and structural design candidates do not support a near term SSTO option. An assisted low Earth orbit capable SSTO may be considered a near term option. Sled and rocket assist design options based on current rocket propulsion capabilities plus at least one Have Region structural design concept could provide the performance capability for a low Earth design reference mission. An assisted SSTO design option is considered to be medium to high risk at this time. TSTO is considered to be the lowest risk near term option. Both current propulsion and structural design options would support a TSTO design option. 29

30 Summary Continued Low earth to orbit design reference missions are considered to be the easiest to achieve and the lowest risk option. A Geo transfer orbit design reference mission is highly efficient to perform a total mission model. A Geo transfer design reference mission model would eliminate both SSTO and assisted SSTO vehicle design candidates. The mission velocity requirement is significantly higher than the mission velocity requirements for low Earth orbit missions. A global range sub orbital first stage with a boost-glideskip mission profile requires significantly less mission velocity (approximately 15% less) compared to a low Earth orbit design reference mission. The required first stage mission velocity for a global range boost-glide-skip profile, would require a usable propellant fraction requirement of approximately

31 Summary Continued A supersonic air-breathing turbojet engine is a potential first stage propulsion system for TSTO vehicle design concepts. However, a Mach 4 first stage would achieve less then 3% of the kinetic energy needed to achieve a low Earth-to-orbit design reference mission. A 97% capable second stage would be difficult even for a hydrogen-oxygen fueled second stage. Higher energy design reference missions may not be achievable. A combined cycle turbo-ram engine could increase staging velocity to Mach 5-6, however, the same conclusion about the Mach 4 turbojet apply to a somewhat lesser degree. A hydrocarbon ram-scramjet is currently being developed that would increase the staging velocity to Mach 6-8. It would be very challenging to have a hydrocarbon ram-scramjet engine for a near term system. Rocket propulsion is the best propulsion choice at this time. 31

32 Summary Continued The Lockheed- Martin Have Region structural design test article did not complete testing. X-ray inspection of the manufactured test article revealed faulty welding. The test article weight was heavier then predicted The test article did, however, successfully demonstrate the manufacturability of a 3 mil thick stainless steel liner for the hydrogen tank. The Lockheed Martin X-33 program did not validate a structural test article. The McDonnell Douglas test article was a scale up of the Isinglass test article that was successfully design, built and tested in the 1960s. The actual weight of the Have Region test article was heaver than predicted. Unfortunately, the load transfer structure failed during the first structural load test. subsequently, the test article blew up. The Boeing Have region test article was successfully designed, built and tested. Some buckling of a lower surface skin panel occurred during reentry testing. Close inspection showed the test article was not built to specifications. The test article contained over 95% of the parts required for the entire vehicle and weighted less then predicted. As of two years ago the test article was still in storage at the Boeing Kent facility. A half scale composite hydrogen tank was successfully designed, built and tested in the NASP program. 32

33 Conclusions A large number of near term reusable launch vehicle design concepts have been investigated since the late 1970s. New design concepts continue to emerge today. While the Boeing RASV was the preferred choice during Science Dawn, it is most likely not the preferred choice today. Missions, operational requirements, payloads, and technologies have continued to evolve. A low Earth orbit design reference mission may not be an appropriate design reference mission. A geo transfer design reference mission should be investigated further. Propulsion choices are limited. Air-breathing propulsion systems do not appear to be an appropriate choice. Current rocket propulsion choice is the SSME. The Rocketdyne linear aerospike engine would require further work and funding to transition from a prototype into a long life, reliable, flight weight operational engine. A Pratt &Whitney expansion cycle is a highly desirable new engine, but would need development. 33

34 Conclusions and Recommendations Continued As a result of Have Region only one validated structural design concept emerged. The Boeing concept has a propellant utilization fraction of 0.88, sufficient for an assisted SSTO low earth orbit, or a TSTO design concept with a Mach 23 staging speed. Conduct an independent in depth design study to verify validity of the near term RLV ( pending NSF proposal, $1 million ) Support the development of the Pratt & Whitney class of high thrust hydrogen oxygen expansion cycle rocket engine as part of the NASA rocket engine development program. A 2008 operational prototype can be done with acceptable risks Validated structural concept(s) Existing materials and manufacturing techniques Existing rocket engines (require modifications) Boost-glide-skip first stage mission profile minimizes performance and operational requirements Substantial margins available Modified existing subsystems and avionics 34