Mitchell Burnside Clapp
clappm@plk.af.mil
Aerial Propellant Transfer to Augment the Performance of Spaceplanes
Captain Mitchell Burnside Clapp Phillips Laboratory, Kirtland AFB, NM 87117
William Nurick, Frank Kirby, Ed Nielsen, Robert O'Leary, and Ray Walsh
W. J. Schafer Associates, Inc., Calabasas, CA 91302
and Daniel P. Raymer Conceptual Research Corp., Sylmar, CA 91392
In-flight propellant transfer to a rocket-powered aircraft permits it to achieve orbit with relatively little propellant compared to taking off fully loaded from a runway. The weight of many key components, such as wings and landing gear, is substantially reduced. The vehicle takes off like a conven tional aircraft under rocket power from two of its seven engines, using jet fuel (JP-5) and a non-cryogenic oxidizer (H2O2). After rendezvous with and propellant transfer from a tanker aircraft, the vehicle lights all its engines, accelerates to high speed, and pulls up into a steady climb into orbit.
Non-cryogenic, non-toxic propellants permit the propellant transfer to use existing tankers, and a small aircraft similar in size to an F-16 could dem onstrate the capability and achieve orbit. Many important military missions could be performed by such an aircraft. The concept is sufficiently simple that relatively little in the way of new facilities or support equipment is required.
The mass of a single stage rocket vehicle at the beginning of its mission Mo is related to the mass at the end of the mission Me by the rocket equation:
Mo/Me = exp(Dv/Isp g)where the symbol Dv represents the required mission velocity change, including losses due to aerodynamic drag, gravity, back pressure on the engines, steering, and so forth, Isp is the specific impulse of the engine in a vacuum, (defined as the number of pounds of thrust per pound per second of mass flow through the engine), and g is the acceleration of gravity at Earth's surface, which appears in the equation to convert mass to force so that the argument of the exponential is dimensionless.
Single stage to orbit flight is technically challenging because the *v necessary to achieve orbit (30,500 ft/s, typically) imposes mass ratios that are difficult to achieve with current structural technology. The usual approach is to seek more energetic propellants with high Isp values. Airbreathing approaches are also an attempt to achieve large Isp. This tends to involve propellants that are not very dense and difficult to handle, such as liquid hydrogen, or impose surpassingly difficult design and operations problems such as those that have afflicted the National Aerospace Plane program.
Single stage rocket vehicles fall into three principal configuration categories: vertical takeoff/horizontal landing, such as the SSTO/R vehicle proposed by the NASA Access to Space Study, vertical takeoff/vertical landing, such as the McDonnell Douglas Delta Clipper design, and horizontal takeoff/horizontal landing, such as the Boeing RASV or British HOTOL designs.
Between the first two of these, there is no obvious distinction in terms of empty weight. Credible design studies appear to give similar weight estimates for similar vehicles. Horizontal takeoff and landing vehicles, however, tend to be much heavier for a given payload because of the unique requirements imposed by runway takeoff. Wing loads at rotation and the weight of landing gear are of particular concern. Generally horizontal takeoff and landing vehicle designs tend not to be pure single stage to orbit, but rely instead on sled launch or auxiliary boosters to reduce gross weight.
The purpose of this paper is to discuss another approach for operating spaceplanes off conventional runways with conventional facilities: using in- flight propellant transfer to reduce the takeoff gross weight of a rocket powered aircraft, and hence its size, weight, and cost. This is not an attempt to solve the rocket equation problem by means of increasing Isp, but by decreasing Dv. Beginning the mission to space from tanker altitude and airspeed reduces the amount of propellant that must be expended overcoming drag and gravity losses. The emphasis is on maximizing the use of existing components and keeping the design as simple as possible. Hence, we will use existing tankers, landing gear, and conventional technology as much as possible and examine the resulting size of the vehicle. A contracted six- week study between Phillips Laboratory, WJ Schafer Associates, and Conceptual Research Corporation developed this concept further. The ground rules for the study were:
Horizontal takeoff like an aircraft
Two engines firing at takeoff
Propellant transfer at 40,000-43,000 ft
Hydrogen peroxide and jet fuel propellants
Power-off landing
LEO mission
Throttling during propellant transfer
Maximize use of existing facilities and support equipment
Conservative design assumptions
Tanker aircraft selection
In-flight refueling is commonplace in the US Air Force and Navy today. Two
systems are used: the Navy's probe and drogue system and the Air Force's
boom
refueling system. The probe and drogue system requires the pilot of the
receiver aircraft to do all the work, and transfers about 250 gallons per
min
ute. The boom system requires some cooperation between the boom operator and
the receiver aircraft pilot, and can transfer 1200 gallons per minute. The
boom refueling system was selected for this design because of its high rate
of propellant transfer.
Two types of tankers use the boom system today: the KC-10 and KC-135. Of these, the KC-135 is smaller, less expensive, and more readily available. Of particular interest is the KC-135Q and KC-135T. These aircraft have an isolated fuel system, from which the tanker's own engines cannot draw. This will allow dedicated rocket propellant tankers to operate with only minor im pact on the tanker's own systems. To avoid a costly development program, and the need to completely re-engineer the transfer system, the propellant carried by the tanker should be non-cryogenic and non-toxic.
There are only a few non-cryogenic oxidizers available: red fuming nitric acid, nitrogen tetroxide, and hydrogen peroxide are the obvious choices. Of these, only hydrogen peroxide is non-toxic. It has other advantages as well. It is very dense (1.432 g/cc in 98% concentration). It has a vapor pressure about one-ninth that of water. It is relatively inexpensive because it is an ordinary industrial chemical rather than a dedicated rocket propellant. Because it is a good coolant, ordinary JP-5 rather than expensive RP-1 can be used as the fuel. Although some special precautions must be taken to pre vent it from decomposing in the presence of impurities, it is a stable mole cule, and once those precautions have been taken it essentially handles like water.
Detailed analysis of a hydrogen peroxide/jet fuel engine indicates the following performance figures at a mass mixture ratio of 7.30:1 (oxidizer:fuel). The two columns in Table 1 are for the two versions of the engine. The first version is operable at sea level and permits the aircraft to take off, rendezvous with the tanker, and transfer propellant. The second version is only operable at tanker altitude or above, and is optimized for the climb to space.
Table 1 Hydrogen peroxide/jet fuel engine performance
Climb Engine Takeoff Engine
Chamber pressure 3000 3000 psia
Exit plane pressure 1.0 5.7 psia
Expansion ratio 240 70 --
Ideal Isp (shifting equilibrium) 354 340 sec
Losses due to:
geometry 2.4 2.4 sec
finite rate chemistry 1.8 1.0 sec
viscous drag 7.8 6.6 sec
energy release efficiency 6.7 7.3 sec
Delivered Isp (in vacuum) 335.3 323.1 sec
Thrust 19930 19210 lb
Weight 280 310 lb
The advantages of the aerial propellant transfer concept are threefold. First, the propellants are at a very high density -- 1.32 g/cc of propellant at the mixture ratio given. This leads to a smaller vehicle and the capability of transferring up to 155,000 pounds of hydrogen peroxide from the tanker to the receiver. Second, they are non-cryogenic, so that the modifications to the KC-135Q or KC-135T model tanker will be minimal. Finally, the mixture ratio is unusually high. At a mixture ratio of 7.30 to 1, 88 per cent of the benefit of aerial propellant transfer is available if one propellant only is transferred. This helps with keeping the operation simple and removes some safety concerns with simultaneous propellant transfer.
The mission profile begins with a takeoff from a conventional runway using the two takeoff rocket engines for thrust. The aircraft is loaded with all the fuel it needs for the climb from the tanker to orbit. It also has fuel and oxidizer aboard sufficient for 15 minutes of atmospheric flight. The total weight of the vehicle at takeoff is about 50,000 pounds, but by the time it achieves tanker rendezvous at 43,000 feet and 0.85 Mach number its weight has dropped to about 38,000 pounds.
When the aircraft meets the tanker it takes on about 147,000 pounds of hydro gen peroxide. It then disconnects from the tanker and climbs to space. As it inserts into orbit, its weight has dropped to about 16,500 pounds. After performing its orbital mission, the aircraft reenters and glides to a normal landing at a runway.
The weight buildup of the vehicle will determine whether it is possible to enclose the required volume of propellant in an aircraft that weighs little enough to permit that propellant to launch it into space. The table below indicates the assumptions for each of the major weight components and the total weight of the system.
The basic assumptions made for the vehicle are to apply conventional structural technology by forming the blended wing/body of the aircraft from ordinary aluminum alloy. The thermal protection system technology deemed suitable for this application is carbon/silica carbide for the nose cap, DuraTABI for acreage areas on the lower surface, and a lightweight blanket insulation for the upper surface. The crew cabin accommodations are austere, as in the U-2 reconnaissance aircraft.
Table 2 Weight Breakdown (pounds)
Structures Group 6,686
Wing 1,572
Vertical tail 739
Fuselage 2,924
Main landing gear 916
Nose landing gear 243
Engine mounts 292
Propulsion Group 3,091
Engines 2,120
Fuel system 971
Equipment Group 1,181
Flight controls 372
Instruments 142
Avionics 567
Furnishings 100
Mission-specific Group 4,000
Reaction controls 400
Life support 800
Thermal protection system 2,800
Total Empty Weight 14,958
Load Group 33,494
Payload 1,000
Crew 440
Propellant 32,054
Takeoff gross weight 48,452
Tanker rendezvous weight 37,380
Oxidizer transfer 146,870
Gross light-off weight 184,250
Unlike most spaceplane designs, this vehicle needs to have a particularly high subsonic lift to drag ratio. This is necessary for two reasons. First, the requirement to fly in the atmosphere on the rocket engine impels the designer to minimize thrust required, so that the rocket propellant load at takeoff remains small. Second, the vehicle's gross weight changes by a factor of about 4.5 during propellant transfer. The maneuver will be very difficult for the pilot to fly if the aircraft does not have a good cruise lift-to-drag ratio.
The aircraft features a highly blended design to maximize volume. The double- delta planform is adopted to provide minimal change of the aerodynamic center over a broad speed range, and also to provide a large strake to hold fuel and oxidizer so that the center of gravity does not move as the propellant is consumed. The overall wing area is 780 square feet. The wing loading is suffi ciently low that no lift devices such as flaps or slats should be needed for takeoff or landing, especially with the enormous thrust available from the rocket engine. Low wing loading may also moderate the thermal environment during reentry.
Unlike most space vehicles, it will be possible to test the aircraft proposed here in a conventional flight test environment. No special range require ments beyond what is conventionally available at, for example, Edwards AFB should be required. Because there are aviators aboard the vehicle, no requirement for a destruct package exists. Aside from storage areas for the new propellant, it should not prove necessary to construct any new facilities for any phase of this program.
The flight test program could begin in a conventional build-up fashion, starting with taxi and ground tests, first flight, performance, and flying qualities testing. This phase of the program would emphasize handling qualities while connected to the tanker boom, because the oxidizer transfer will quadruple the weight of the aircraft when it takes place. Once the flight control system has been qualified, transfer of steadily increasing amounts of oxidizer would support envelope expansion and flight to increased altitudes and airspeeds. Exoatmospheric flight and reentry could be investigated, and the operational envelope of the thermal protection system could be determined. The capability of the system to perform ballistic transfers to anywhere on earth within one hour could be demonstrated.
Loading the aircraft with fuel and oxidizer at 7.30:1, up to the maximum takeoff weight, could also permit exoatmospheric flight without propellant transfer. The ballistic ferry range of the aircraft under these conditions is about 3200 nautical miles, allowing for some aerodynamic range extension at the end of the trajectory.
An orbital flight attempt would follow the envelope expansion phase. Investigation of on-orbit flying qualities could proceed at this point, as well as an experimental determination of reentry cross range. One sub-phase of the orbital flight test program of particular interest would be on-orbit propellant transfer. If the aircraft were completely refueled in low earth orbit, it would have enough Dv to visit anywhere in cislunar space, such as geostationary orbits, or to perform multiple plane changes and visit many different points on a single mission. Reentry from increased altitudes and entry speeds could be tested, yielding an assessment of the capability of a high temperature reentry capability in realistic conditions. Conclusions Using in flight propellant transfer to reduce the Dv needed to fly to space makes it possible for a fighter-sized aircraft to achieve orbit. The enabling technology to do this is non-cryogenic, non-toxic rocket propulsion based on H2O2 and JP-5. Developing this capability permits a variety of militarily significant capabilities to be demonstrated.