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White Paper 03 April 2006

THOUGHTS ON THE POSSIBLE EXISTENCE OF THE

"BLACKSTAR"

O RB IT AL SP ACE P L AN E S YS T E M

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1200 Ashwood Parkway, Suite 506 | Atlanta, GA | USA 30338 1+770.379.8000 | 1+770.379.8001 [Fax] | www.sei.aero | [email protected]

G OAL S O F T H IS R EV IEW The purported reusable, crewed two-stage-to-orbit flight system referred to as "Blackstar" is briefly technically assessed in the context of previous studies and technological developments. Leading features of its "SR-3" first-stage carrier aircraft and its "XOV" upper-stage spaceplane vehicle are commented upon. Several key design problem areas are pointed out, e.g., takeoff/landing gear, transonic acceleration, upper-stage propulsion. Finally, an attempt is made to positively envision the Blackstar vehicle system as being a technically feasible development, an exercise leading to the still unanswered question: did it exist?

References: 1. 2. Scott, W.B., Aviation Week & Space Technology, 6 March 2006, pp. 48-53 (three articles) Jenkins, D.R. and Landis, T., North American XB-70A Valkyrie, Specialty Press, 2002

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B AC KGR O UN D The following is a summary of what is known about the Blackstar system based upon an article in the March 2006 issue of Aviation Week & Space Technology and taken from Wikipedia [www.wikipedia.org]. Source: Wikipedia [http://en.wikipedia.org/wiki/Blackstar_(spaceplane)] Blackstar is the reported codename of a secret United States orbital spaceplane system. The possible existence of the Blackstar program was reported in March 2006 by Aviation Week & Space Technology (Aviation Week) magazine; the magazine reported that the program had been underway since at least the early 1990s, and that the impetus for Blackstar was to allow the United States Government to retain orbital reconnaissance capabilities jeopardized following the 1986 Challenger disaster. The article also said that the United States Air Force's Space Command were unaware of Blackstar, suggesting it was operated by an intelligence agency such as the National Reconnaissance Office. Aviation Week speculated that such a spacecraft could also have offensive military capabilities (a concept colloquially known as "The Space Bomber"). The magazine also said that it was likely that Blackstar would be mothballed, although it is unclear whether this is due to cost or failure of the program. Aviation Week describes Blackstar as a two stage to orbit system, comprising a high-speed jet "mothership" aircraft (which Aviation Week referred to as the SR-3). Its description of SR-3 is similar to the North American B-70 Valkyrie Mach 3 strategic bomber and to

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patents filed in the 1980s by Boeing. The SR-3 would carry a second, smaller airframe, codenamed the XOV (eXperimental Orbital Vehicle). This rocket-powered spaceplane, with similarities to the X-20 Dyna-Soar project, would be released by its mothership at an altitude of around 100,000 feet. The XOV would then light its rocket motor and could achieve both suborbital and orbital flight; one source quoted by Aviation Week estimates the XOV could reach an orbit of 300 miles above the Earth, depending on payload and mission profile. The XOV would then reenter the atmosphere, fly like a normal aircraft (possibly using aerospike engines, similar to those used by the Lockheed Martin X-33), and would land horizontally on a conventional runway. This combination of jet-powered mothership and a smaller rocket-powered spaceplane resembles the civilian Tier One spaceplane system, but capable of much higher velocities and of thus attaining orbit.

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GENE RAL DE SC RIPT ION The "Blackstar" two-stage to orbit launch vehicle system consists of an "XB-70 like" large airbreathing engine powered supersonic aircraft boost stage, carrying an underslung "boost-glide" fashioned upper stage vehicle1. The boost aircraft is referred to as the SR-3 and the upper stage as a "spaceplane" or an XOV (for experimental orbital vehicle). The SR-3 aircraft features a clipped delta wing with vertical tails at its extremity, and an extended fuselage. Shown as the powerplant is a set of supersonic air intakes, an extended flowpath and a set of either six, or four exhaust ports. The vehicle is described as having a laterally separated pair of flowpaths terminating with two exhaust ports. Given space requirements for the "semi-conformal" housing of the upper stage along the aircraft's ventral surface, this approach seems appropriate. The XOV upper stage vehicle is shown to be configured as a blended wing-body with its outer wing panels canted slightly downward. A large bluntly terminated aft end is shown to be outfitted with four rectangular engine exhaust ports. Its large stub dorsal vertical fin is said to serve as the structural pylon connection to the boost vehicle. Both stages are implied to be crewed (cockpits with high-speed windscreens are shown).

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B AC KGR O UN D O F S I MIL AR ( OR REL ATED) T STO SY STE M C ONCEPT S Two-stage reusable orbital launch system concepts have been extensively documented since the 1950s beginning with the original USAF-sponsored aerospaceplane study effort, and continuing with the NASA (MSFC-FPO) Reusable orbital Transport (ROT) study of the early 1960s. Following these all-rocket were powered developed, concepts, the combined stages airbreathing/rocket systems upper

continuing as all-rocket systems. An early set of examples of these combined-propulsion systems were the Lockheed-designed TSTO HTHL vehicles expressed in the NASA-sponsored mid-1960s study under Contract NAS7-377 led by The Marquardt Corporation. Moving to contemporary times, the German Saenger project espoused a TSTO system with its boost stage powered by a set of hydrogen-fueled turbine/ramjet engines, and its upper stages (both reusable and expendable types) being all-rocket powered. A more recent set of TSTO concepts that used NASA's revolutionary turbine accelerator (RTA) firststage engines was developed by Boeing (FAAST) and Lockheed Martin. The Boeing concept staged at a relatively low speed (at RTA termination) requiring a combined airbreathing/rocket powered upper stage. The Lockheed Martin concept, using a ramjet-extended higher staging speed had an all-rocket upper stage, and, in this respect was similar to Saenger.

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C O MMENT S ON T HE PRO JE CTED SR-3 BOOST AIRCRAFT By implications of being "XB-70 like" the Blackstar system staging speed would presumably be in accord with that airplane's Mach 3 top-speed capability. However, since the SR-3 vehicle is, by appearances, not required to sustain a thermally-hostile cruise condition, being strictly an acceleration means, it might be capable of a somewhat higher staging speed which would be advantageous in terms of the post-staging, high delta-V requirements faced by the upper stage in going to orbit. If the SR-3 vehicle was to use the XB-70A's J-93 class afterburning turbojet engines, which were rated for continuous afterburner-operation at Mach 3, and if these were to be then modified for water injection (and possibly liquid oxygen injection) along the lines of the MIPCC effort, a transient capability to accelerate to, say, Mach 4 ­ 5 might result. MIPCC stands for mass injection and pre-compression cooling. Experimental investigations of this interesting approach are currently underway. Configurationally, the clipped delta-wing layout with tip-mounted vertical fins would likely negate the unique variable-geometry "drooped" outer wing panels of the XB-70. These were used to gain high-speed compression lift and for lateral-stability augmentation, important steps for improving the aircraft's cruising-flight performance. Not having this feature in the non-cruise SR-3 might not pose a significant disadvantage, and this would reduce both physical and operational complexity of this machine, vis-à-vis the XB-70A aircraft.

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The question arises as to the overall takeoff grow weight (TOGW) of the SR-3/XOV combined-element vehicle. This will be gone into later, but as a starting-point reference, the related characteristics of the XB-70A aircraft can be instructive. The stated maximum weight of this airplane was 540,000 lbm, its rated takeoff and landing weights being somewhat below this figure. It was powered by six General Electric YJ93-GE-3 engines, each providing a maximum sea-level thrust in full afterburner of 28,800 lbf. Based on an allowable TOGW of 520,000 lbm and a total thrust of 172,800 lbf, the overall thrust/weight ratio was 0.33, i.e., thrust was one-third weight. But the SR-3's engine count might be less than the six engines of the XB-70: "At least four engine exhaust ports are grouped as two wellseparated banks, with ports on each side of the aircraft's centerline"1. Having but four engines would suggest that, assuming that the SR3/XOV combination was at least as heavy as the XB-70, higher thrust engines would be needed. The only turbojet engine on record that measurably exceeded the J93 engine's thrust in the late 1950s was the developmental General Electric supersonic transport engine, the GE-4 afterburning turbojet. This engine was initially rated at 52,600 lbf, but "A later version of the engine achieved 63,200 lbf on 19 September 1968, establishing it as the most powerful jet engine in the world at that time"2. Just as a "thought exercise," suppose that the GE-4 had been developed to operational status (in fact, the U.S. SST program was to be terminated, as was this engine program). Assuming that the SR-3 aircraft was now to be powered by four GE-4 engines providing a total thrust of 252,800 lbf,

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at the overall thrust/weight ratio of the XB-70 (0.332), the SR-3/XOV TOGW would then be as high as 761,500 lbm. This figure approaches B747-class aircraft weight values ­ and is almost 50% more than the TOGW of the XB-70. T R ANSON I C ACCELE RAT I ON P ROBLEMS Turbine-powered supersonic aircraft generally have an inherent problem in the transonic flight regime, where aerodynamic drag forces rise sharply over subsonic values without means of measurably increasing thrust in compensation. The resulting reduction in thrust-minus-drag (TD) can greatly extend the aircraft's transonic "push through" time causing unwanted increased fuel usage. And, if horizontal-flight T-D goes to zero, unless you dive, you won't go supersonic. In the case of the Blackstar system, even if the XOV upper stage was to be extensively conformally integrated with its SR-3 carrier vehicle, unless it were to be completely encapsulated in the boost vehicle's fuselage, it would significantly increase the combined vehicle's drag characteristics, over those of the SR-3 alone. The resulting impact on the system's turbine-powered transonic flight performance would be of critical importance. Presuming an effective drag-minimization design effort was to be mounted, the only avenue remaining is to increase vehicle thrust.

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Looking to the main engines, already at full afterburner, unfortunately the MIPCC approach for thrust augmentation has been found not to be very effective in the transonic-flight regime. The remaining remedial possibility is to bring on auxiliary propulsion, the most obvious means being rocket propulsion, perhaps using the propulsion system of the upper stage. But this would entail a significant added-mass penalty and lead to both physical and operational complexity. We will return to this option below. TAKEOFF/L ANDING GEAR E XTENSION NE E DS Referring back to the XB-70A developmental experience, its high-speed takeoff (rated at 205 knots) and landing gear proved to be a considerable operational challenge, with a number of gear failures experienced. To fit into the limited retracted-gear storage volume in the fuselage, the fourwheel main gear went through a complicated folding and turning process prior to being positioned "in the well." This is likely to be the case with the SR-3 main gear as well. But, in addition, because of the presence of the stowed upper stage with a significant portion of its bulk protruding below the SR-3's lower mold line, the gear will have to be extended beyond its normal length, further complicating this major set of retractable assemblies. Also, it seems probable that the spacing of the two main gear elements would have to be extended beyond the design-optimal spacing in order to accommodate the upper stage with adequate clearance. These forced design changes in an already complex gear design (ala XB-70) will tend

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to add additional inert weight, as well as reliability-compromising complexity. U P P E R S T AGE P RO P ULS I ON The descriptions provided in the articles is confusing in that it calls out "linear aerospike" engines and shows four rectangular exhaust ports1. But elsewhere it notes the presence of vehicle-underside "strakelets" that direct airflow into engine inlets, adding that the airflow might then be processed through a complex system of composite-material ducts, etc. Noting that the linear aerospike design applies to liquid-propellant rocket engines, with this "controlled airflow" reference unrelated to rocket propulsion, does this suggest that both high-speed airbreathing and rocket engines are to be installed in the orbiter vehicle? And are the rocket engines configured as linear aerospike systems, or would the more conventional bell-nozzle type be used? We will return to these key upper stage propulsion system questions below. B ORON -B ASE HIG H ENERGY FUEL S A "fuel breakthrough" is noted to have greatly accelerated the spaceplane's propulsion system development1. A boron-based gel of toothpaste consistency is claimed to have compact, high-energy content characteristics, and was provisionally assumed to be used in this system. There is an extensive history of work on high-energy fuels (HEF) by the U.S. Department of Defense (DoD) in the mid-/late-1950s (see Appendix C of Reference 2)2. A family of "Alkylborane" fuels were placed under

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exploratory development for high-speed aircraft engines, both turbine and ramjet types as reported here. HEF-3, ethyldecaborane, was nominally specified for the XB-70 program as a means of extending the bomber's range. This fuel had a heating value of 25,000 Btu/lbm, as compared to 18,000 for regular JP fuels. When used just in the airplane's afterburners a range extension of about 15-percent seemed achievable. The J93-GE-5 engine was to be developed for this HEF-afterburner operation. However, it was soon determined that the use of HEF fuels introduced both material compatibility and operational problems requiring additional development work. But, in mid 1957, "the Pentagon unexpectedly reduced the overall scope of the high energy fuel endeavor and canceled the J93-GE-5 engine program"2. This decision effectively removed the HEF components of the XB-70 program. Interestingly, just prior to this termination decision, Air Force management took steps to shift the HEF-program application priorities from the B-70 to the Bomarc ramjet-powered interceptor missile program. Its Marquardt RJ-43 ramjet engine was flight tested on an X-7 vehicle using HEF. Its exhaust trail, normally invisible, was a prominent and persistent bright white, rope-like streak trailing behind this supersonic flight vehicle. However, from this reviewer's perspective, there is no open-literature record of note covering the use of HEF as a rocket propellant.

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ENVISION ING THE B LACKSTAR SYSTE M FROM THE PERSPECT IVE OF THE PRECE DING REVIEW SR -3 F IRST STAGE C ARR IER AIRC RAF T On a hypothetical basis it is possible to technically envision the supposed Blackstar system, along the lines of that described1. An effort to do so will now follow, keeping in mind the points made above. If the SR-3 first-stage vehicle used four GE-4 SST type engines, as conjectured earlier, and weighed in at B747 levels of takeoff mass, it would seem to be capable of carrying a considerable XOV second-stage spaceplane mass to a near-Mach 3 staging condition at high altitude. Takeoff noise levels would be quite high1. In keeping with the implied sustained afterburning turbojet operation, the dynamic pressure at this terminal point would be at significant levels (i.e., a non-vacuum staging would follow). As conjectured, the four SST engines would be installed in two wellseparated flowpaths, each equipped with a supersonic, probably variable-geometry inlet1. This spacing arrangement, in the direction of that used in the Concorde SST, would be necessary in order to centrally house the "conformally integrated" underslung upper stage. A conventional retractable tricycle takeoff/landing gear, one capable of high-speed takeoff operations, would be installed. However, it would likely have to be considerably extended in length to accommodate the upper stage volume protruding below the SR-3 lower mold line. Also, the

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main gear would likely be spaced apart more than normally to clear the upper stage wing-body configured vehicle. One is reminded of the extended-gear used in the Convair B-58 Hustler aircraft with its very large underslung pod assembly. Generally, the SR-3 takeoff/landing gear would probably be both heavy and complex. TR ANSONIC PU NC H -THROU GH CAPAB IL ITY Responding to the characteristic transonic drag-rise problem discussed earlier, one in this case exacerbated by the parasitic high-drag profile of the blunt-ended upper stage as envisioned1. Even with extensive dragreduction aerodynamic fairings, if used, it is believed that added rocket thrust would have to be employed to efficiently "punch through" the transonic flight regime for continued supersonic acceleration. This might be where the upper stage's two upper-stage outboard solid-propellant rocket units mentioned would come in. Once through transonic, with these rockets exhausted, to reduce the inert weight being further accelerated, they could be ejected aft-wise, and perhaps safely recovered by parachute means like the Shuttle SRBs. Otherwise, they might constitute a collateral-damage hazard, depending on the overflight pattern being followed. O N TO STAGING CO NDITIO NS The afterburning turbojet powered vehicle would continue to climb and accelerate to its staging condition of Mach ~3, or possibly somewhat higher. The use of the MIPCC-type water-injection process was noted

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earlier, as a means for both thrust augmentation and for extending an engine's upper speed limit. At the staging point, the upper stage would be literally "dropped" from the SR-3 carrier aircraft which would momentarily climb as it was relieved of this large mass. The XOV spaceplane would then start its propulsion system in a presumed ramjet mode, possibly using HEF-3 as had been developed under the DoD program as described earlier. Recall, the Bomarc RJ43 ramjet engine was flight-tested on HEF. But this surmise can be questioned in view of the logistics and vehicle operational complexities involved with HEF, as opposed to conventional JP hydrocarbon fuel as used in the SR-3, and possibly to be used as the fuel for the upper stage's follow-on rocket operation post ramjet mode. In fact, the use of airbreathing propulsion in any form by the XOV stage can be questioned in view of its relatively small delta-V contribution before the rocket system would be engaged to accelerate the vehicle to much higher orbital-insertion speeds. CO NVE NT I O N A L RO CKE T E NG I NE S L IKE L Y U SE D While the use of linear aerospike engines are suggested, the need for this unconventional rocket engine type can be questioned1. For one thing, with its launch at high altitudes and low atmospheric static pressure level, there is no need for the unique "altitude-compensation" feature of the aerospike design. For another, the development experience with the aerospike engine has been almost strictly on hydrogen/oxygen propellants. In view of its relatively compact configuration, it is not

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considered that this high-energy, but bulky propellant combination would be used here. From these considerations it would seem that conventional storablepropellant bell-nozzle rocket engines would be used to power the XOV stage over the significant delta-V span required beyond its relatively low staging speed. Either hypergolic storable propellants, e.g., NTO/MMH, as used in the Shuttle OMS and RCS systems, or hydrogen peroxide/JP propellants would likely be used. The latter choice was recently made in designing the Air Force/SpaceWorks Quicksat TSTO system. The use of such non-cryogens would avoid the possible complications of an SR-3 liquid oxygen topping system, had that cryogenic oxidizer been used. A T H I GH -SP EE D FL I GH T CO ND I T ION S Rocket-propelled acceleration of the XOV upper stage to either a high sub-orbital atmospheric flight condition, yielding a classical boost-glide type mission profile, or to a minimal orbital-insertion in-space condition, would then follow. The vehicle would then carry out its designated mission operation in a once-around flight back to base, or after a few low orbital passes. It would then be prepared for its descent and landing phase. It would then reenter, if in space, and decelerate aerodynamically and complete a controlled aerodynamic-flight descent to its assigned landing base. An unpowered horizontal glide-in landing would then be performed. If a conventional runway, rather than a dry lake bed as used by the X-15 at EAFB, were to be used here, the XOV's purported centralbody mounted skid might not be welcomed in view of runway damage

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that might ensue1. The option of a installing a high-speed main landing gear is a consideration here RETURN TO BASE Post staging, the large SR-3 aircraft would decelerate, descend and conduct a conventional horizontal rolling-gear landing. It could then be prepared for its next flight and remating with another (or the same) XOV stage. The XOV is incapable of low-speed self-ferry operation. If it landed at a location other than a Blackstar operating base, the XOV would require separate transportation to its main base. The use of specially modified C-5A Galaxy transports for this relocation process was cited1. But less sophisticated means might also serve, e.g., the X-15 was transported over-the-road in a special tractor/trailer rig.

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CONCL UDING REMARKS So, in this highly speculative concluding section of this review of Blackstar system possibilities, it can be surmised that, from an extended aerospace engineering standpoint, such a crewed reusable two-stage-toorbit vehicle system might be considered a "doable proposition." The main issue of whether this system was actually developed and flown is another question -- at this point, one that will have to remain unanswered.

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ABB REVI AT ION S AN D ACR ON YMS USE D

DoD EAFB HEF HTHL MIPCC OMS RCS ROT RTA SST T­D TOGW TSTO U.S. Department of Defense Edwards Air Force Base (U.S. Air Force) High-energy Fuel Horizontal Takeoff and Landing Mass injection and pre-compression cooling Orbital Maneuvering System Reaction Control System Reusable Orbital Transport Revolutionary Turbine Accelerator Supersonic Transport Thrust-minus-Drag Takeoff Gross Weight Two-stage-to-orbit

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ABOUT THE AUTHOR

Mr. William J.D. Escher Senior Technical Fellow SpaceWorks Engineering, Inc. (SEI) Mr. William J.D. Escher is Senior Technical Fellow at SpaceWorks Engineering, Inc. (SEI). Mr. Escher has 50 years experience working in the aerospace industry divided among nearly equal tenures in the government, industry, and selfemployed arenas. He has had a long-term involvement with space transportation and related propulsion developments that stretch back to the United States' first attempts to place scientific satellites in orbit during Project Vanguard run by the Naval Research Laboratory. Mr. Escher has also held positions at multiple NASA field centers as well as at NASA Headquarters, Rocketdyne, The Marquardt Corporation, Kaiser-Marquardt, and Astronautics Corporation of American. Additionally, Mr. Escher spent fifteen years as a self-employed consultant in the areas of terrestrial energy systems and environmental betterment technology. Just prior to joining SEI, Mr. Escher served as a senior systems engineer in the Applied Technologies Group of Science Applications International Corp. (SAIC) in Huntsville, Alabama. At SAIC, he applied his expertise in airbreathing/rocketbased combined-cycle propulsion to the development of propulsion technology databases and advanced space transportation system assessments. Mr. Escher holds a Bachelor of Science in Engineering degree from The George Washington University, having earlier undertaken mechanical engineering studies at Cornell University and Cleveland State University. At Cornell, he served as the president and experimental committee chairman of the Cornell Rocket Society. He has more recently conducted graduate studies at the University of Southern California and the University of Wisconsin-Madison. He is an associate fellow of the American Institute of Aeronautics and Astronautics (AIAA), a long-term member of SAE International, and a founding participant in the Space Propulsion Synergy Team (SPST).

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ABOUT THE FIRM

SpaceWorks Engineering, Inc. (SEI) is a small aerospace engineering and consulting firm located just outside of metro Atlanta. The firm specializes in providing quick and unbiased analysis of advanced space concepts ranging from space launch vehicles to deep space missions. Five practice areas define the firm: Space Systems Analysis, Future Market Assessments, Financial Engineering, Policy and Media Consultation, and Technology Prioritization. SEI also develops and markets various engineering software tools including specialized disciplinary analysis tools for space vehicle design. The firm's customers include NASA, the U.S. Air Force, DARPA, and various large and small commercial aerospace companies. Engineering services are available on an hourly or fixed-price basis. The firm's capabilities include conceptual and preliminary level modeling of a broad range of future space transportation and infrastructure concepts. Typical systems architectures include: 2nd/ 3rd / 4th generation single-stage and twostage reusable launch vehicle designs (all-rocket, airbreathing, and combinedcycle propulsion options), partially expendable small satellite launchers, heavy lift launch vehicles (HLLVs), military space plane (MSP) concepts, global-reach and strike systems, launch assist systems, in-space transfer vehicles and upper stages, orbital maneuvering vehicles, planetary spacecraft, in-space transportation nodes, propellant depots, and interstellar missions. For more information including a firm overview presentation, technical papers, biographies, etc. please see: www.sei.aero. For media inquires please contact: Dr. John E. Bradford | President | [email protected] | 1+770.379.8007

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1200 Ashwood Parkway, Suite 506 | Atlanta, GA | USA 30338 1+770.379.8000 | 1+770.379.8001 [Fax] | www.sei.aero | [email protected]

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