Recent work at the Phillips Laboratory in Albuquerque, NM has focused on a space access concept involving in-flight propellant transfer. The concept is similar in terms of Delta-V to air launch concepts that have been previously proposed, but has some advantages over these concepts in terms of performance, scalability, safety, and operational simplicity.
The mission of a typical inflight propellant transfer spaceplane begins on a runway with a vehicle loaded with all its fuel but just enough oxidizer to reach the tanker. The aircraft flies to the tanker and under the control of its pilot transfers the oxidizer needed to burn the remaining fuel and climb to orbit. After the transfer, the pilot maneuvers the aircraft onto a trajectory that achieves a low earth orbit, performs whatever mission is required, and reenters the atmosphere to land without engine power at any suitable runway.
Unlike many other spaceplane designs, the requirement for in-flight propellant transfer imposes a corresponding requirement for a high degree of aerodynamic quality. The choice of propellants is a key issue, with O2/RP-1 or H2O2/JP-5 being the best candidates. Altitude and speed at tanker disconnect, propellant transfer rate, thrust-to-weight ratio, and wing loading all interact in complicated ways.
A flight test program for an aircraft of this type would have much more in common with the test programs of conventional aircraft than with conventional rockets. This is partially due to the presence of a pilot on the vehicle who is actually in control of the vehicle, and partially because the nature of in-flight propellant transfer permits a gradual build-up approach to risk management and envelope expansion.
At the moment, to the limits of accuracy of the current work in
this area, the best choice for propellants is H2O2 and JP-5 by a
narrow margin over O2 and RP-1. Hydrogen peroxide and jet fuel are
non-cryogenic and non-toxic. The high mixture ratio of these
propellants makes the required takeoff weight of the aircraft lower,
and the high density makes for a compact vehicle. These effects help
offset the lower performance of this propellant combination. However,
because there are no existing high quality hydrogen peroxide/jet fuel
engines, a O2/RP-1 based design may be preferred for cost reasons.
Air launch vs. In-flight propellant transfer
There has been a great deal of work done on examining space access concepts that are based on air launch as a means of reducing Delta-V to orbit to a level that reduces the performance requirements on a launch vehicle. Essentially, subsonic staging provides benefits over single stage to orbit in several areas. The gravity losses are reduced because the vehicle is in horizontal flight, supporting itself by aerodynamic means rather than by engine thrust. The drag losses are reduced as well. Above 18,000 feet, over half the atmosphere is beneath you. The back pressure losses on the engines are also reduced because the limit for no separated nozzle flow permits larger expansion engines. Finally, there is the possibility of staging above weather, an operational advantage.
Relatively little work has examined the possibility of inflight propellant transfer as an alternative to air launch, however. 1 The reasons for this are unclear. The inflight propellant transfer concept does offer five distinct advantages over air launch.
First of all, the experience base in military aviation with inflight propellant transfer is enormously greater than that for air launch. Perhaps four hundred manned aircraft have ever been released from beneath other aircraft. 2 A similar number of inflight propellant transfers are performed each day. And the number of stores in excess of 50,000 pounds that have been released from aircraft total a few dozen at most. Every modern military aircraft can be refueled in flight, and for many missions it is critical. "Take off, top off and continue with the primary mission" is an everyday operation in the US Air Force.
Second, the separation of two large objects in flight is an inherently risky maneuver. Stores certification history for military aircraft is full of examples of released objects striking the parent aircraft and causing major damage. The risk can be minimized by a number of means, including captive carry testing, wind tunnel work, and build-up flight test, but at some point the certification program must commit to releasing the object -- an all-or-nothing affair. This level of risk can be managed, but doing so drives costs up. Propellant transfer, on the other hand, can be certified by slowly and incrementally flying formation, then near the tanker, then in dry contact with the tanker, then with increasing amounts of propellant transfer, opening the envelope in a very gradual fashion. Inflight refueling accidents are unheard of in flight test. It is a safer activity to certify.
Third, the performance of an air-launch system is subject to some important limitations. Because two airframes are under the influence of one set of engines, the aircraft cannot climb quite as high for separation as an inflight propellant transfer concept can. The interference drag between the two airframes also limits the envelope of the ensemble to some degree. The effect is not enormous, adding up to an advantage of perhaps 250 ft/sec of Delta-V to the inflight propellant transfer concept, but it is noticeable.
Fourth, the inflight propellant transfer concept offers some important advantages in flexibility. The orbital aircraft has the capability to fly suborbital missions without propellant transfer, to distances of 3,000 to 6,000 nautical miles, depending on the aerodynamic configuration. This capability exists because the airframe is capable of independent takeoff and landing. This offers a transcontinental range for a number of alternate missions that are difficult to imagine for an air launch concept. For the same reason, the tanker and orbital aircraft may be based at different locations, and interfaced only in flight. This offers more basing flexibility and removes the requirements for specialized facilities and ground support equipment such as that needed to mate the Shuttle orbiter to its carrier aircraft.
Finally, the carrier aircraft for an air launch concept must be
either an entirely new aircraft or a major, airworthiness affecting
structural modification to an existing aircraft (unless the gross
weight of the orbital segment is very small). The carrier must bear
not only the weight of the propellant, but also the empty weight of
the aircraft as well as its payload. The orbital aircraft must be a
great deal smaller than its carrier for an air-launch concept, while
it can be larger than the tanker (or tankers) for an inflight
propellant transfer concept. This drives the designer to very large
carrier aircraft, which can be expensive.
In order to examine the utility of the inflight propellant transfer concept, and its application to military requirements, a contracted six-week study 3 between Phillips Laboratory, WJ Schafer Associates, and Conceptual Research Corporation developed this concept further. The aircraft that emerged from this study was called 'Black Horse.' The overall intent was to examine the engineering issues associated with an experimental aircraft program which nonetheless retained some residual military utility. Accordingly, the KC-135Q tanker aircraft was selected arbitrarily, and this selection constrains the size of the available propellant and hence the payload of the vehicle. It must be stressed that the Black Horse study was not undertaken to provide a launcher for any particular class of payloads, but to examine a particular technology for military utility.
The ground rules for the study were:
There are only a few non-cryogenic oxidizers available 4 : 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 5 . 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 prevent it from decomposing in the presence of impurities, it is a stable molecule, and once those precautions have been taken it essentially handles like water.
Detailed analysis 6 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 Takeoff Engine Engine Chamber pressure 3000 3000 psia Exit plane pressure 1.0 5.7 psia Expansion ratio 240 70 -- Ideal Isp (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 effic 6.7 7.3 sec Delivered Isp (in vac) 335.3 323.1 sec Thrust 19930 19210 lb Weight 310 280 lb
The advantages of hydrogen peroxide for 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 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.
It is also possible to consider LO2 as an alternative oxidant for inflight propellant transfer. Although it is difficult to get the vehicle as compact and lightweight, and the problems of handling cryogens are not trivial, LO2 has some advantages. It is extremely inexpensive. There are many high quality engines 7 available using LO2 and RP-1, such as the NK-31, D-58, and RD-120 engines, all made in Russia. The experience base for using LO2 is current and intact. If engine development costs are a significant share of the program's total budget after a detailed cost estimate is performed, then LO2 may be competitive as an inflight propellant transfer oxidant with H2O2.
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 hydrogen 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. Table 2 indicates the assumptions for each of the major weight components and the total system weight.
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 prot 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
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 8 . The thermal protection system technology deemed suitable for this application is carbon/silica carbide for the nose cap, TABI for acreage areas on the lower surface, and a lightweight blanket insulation for the upper surface 9 . The crew cabin accommodations are austere, as in the U-2 reconnaissance aircraft.
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 condition that sizes the engine is the thrust required at tanker disconnect, which depends directly on the gross weight and lift to drag ratio at tanker disconnect. The wing area is chosen to provide enough lift to support the gross light-off weight of the vehicle with the smallest area possible
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 lift coefficient at tanker disconnect is about 1.2, for an altitude of 43,000 feet and 0.85 Mach number. The wing loading is sufficiently 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 requirements 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 a mixture ratio of 7.30:1 by weight, up to the maximum takeoff weight, could also permit exoatmospheric flight without propellant transfer. The aircraft has an unrefueled Delta-V of about 14,000 ft/sec, permitting a suborbital ferry range of the aircraft of about 4,800 nautical miles, allowing for some aerodynamic range extension at the end of the trajectory, with a hypersonic L/D ratio of 3.2. 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 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 Delta-V to visit anywhere in cislunar space, such as geostationary orbits, or to perform many plane changes and visit several 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.
Criticisms of Black Horse
In the year since the initial design work for the Black Horse inflight propellant transfer spaceplane was performed, several organizations have raised concerns about the viability of the concept. These focus in many areas, depending on the background and expertise of the reviewer. The purpose of this section is to summarize the criticisms of the basic idea, and the purpose of the following section is to show how they have affected the design of the basic aircraft.
The claimed structural weight of the aircraft, exclusive of landing gear, is 5527 pounds, according to the figures presented in Table 1 . The overall wetted area of the aerodynamic configuration is 2120 ft2. This leads to a weight per unit area of 2.6 pounds per square foot, a common first-cut metric of aerostructure weight. The F-16, a lightweight fighter of similar size also built primarily of aluminum alloy, has an aerostructure weight of 4 pounds per square foot.
No explicit structural margins are presented in the estimates, and adding these could cause the weight to grow further.
The aerodynamic quality of the aircraft appears to be insufficient to fly safely on the tanker boom at the separation conditions stated (43,000 ft, 0.85 Mach). The fully loaded aircraft weighs about 185,000 pounds and must maintain level flight with a wing area of 780 ft2, imposing a lift coefficient of 1.2.
The high lift coefficient at tanker separation presents several severe problems. First of all, the corresponding lift to drag ratio is about 2.1, requiring much more thrust than stated to maintain position on the tanker. A wing as highly swept as the one described in the initial concept is likely to experience wing rock at these conditions. The tanker is itself much higher than it routinely flies for refueling operations, and the angle of attack needed to stay on the tanker could lead to the boom striking the fuselage. There is little margin for pull-up after the aircraft leaves the tanker, complicating trajectory planning. Finally, the weight changes by a factor of more than four from hookup to disconnect, an unprecedented amount.
Like any space access concept, engine Isp is critical to success. The quoted performance of 323 sec for the takeoff engine and 335 for the climb engine is well in excess of any demonstrated H2O2/JP-5 engine. The combination of chamber pressure, closed cycle, and 98% strength peroxide stated in table 1 will probably yield the stated performance. However, several concerns linger over the practicality of the engine used in this concept.
The chamber pressure stated is 3000 psi, near that of the space shuttle main engine. High chamber pressure has driven enormous maintenance and operability problems with the SSME, and developing a robust and operable engine that must operate for even longer than the SSME is a daunting and expensive task. The heat flux at the throat is moderated by an oxygen rich layer of gas at the wall, which robs performance and may cause some serious materials compatibility concerns. Also, the still-high heat flux levels at the throat appear to require the use of a copper alloy material for the coolant passages, and the compatibility of hot hydrogen peroxide liquid with copper is marginal. The turbine inlet is over 1800 F, and is very oxygenated as well, being driven by decomposed peroxide gas. Neither condition is a problem in itself, but the combination can be extremely challenging from a materials and operability standpoint.
The aircraft as described can just barely place a 1000 pound payload into a 100 x 50 nmi orbit at 28.5 of inclination. The change in Delta-V needed to get to a 100 nmi circular polar orbit will consume all the available payload and then some, leaving the entire concept of marginal utility.
The spaceplane's thermal protection system was designed to protect it during a reentry trajectory modeled after the HL-20's. The heat fluxes that result from that assumption may require aggressive active cooling of leading edges, and with no cryogens aboard the aircraft it is not clear how this can be accomplished.
consumption during transfer
The aircraft is consuming propellant at a significant fraction of offload rate by tanker disconnect. This means that the net offload of the tanker at the moment of disconnect is much less than the capacity of the tanker, leaving the aircraft unable to achieve the gross weight needed to achieve orbit. Also, the fuel needed to burn with the oxidizer consumed during transfer has not been accounted for adequately.
High-concentration hydrogen peroxide handling
Because hydrogen peroxide is unstable and can decompose exothermically, the least accident can lead to a catastrophic explosion. There is no current source for the 98% hydrogen peroxide required for the concept, and the costs of handling and storing the new oxidizer could be prohibitive. Also, the notion of transferring peroxide between two aircraft in flight is certain to introduce contamination. Such contamination could lead to a potentially fatal mishap during ascent.
Additional analysis to address these criticisms
The structural weight estimates were developed using statistical weight estimation methodology 10 . Margins are included in the individual weight estimates, since the methodology predicts the actual weights of existing aircraft that have already suffered the weight growth during design. When the parameters that have a strong effect on wing and fuselage weight (wing loading, wing thickness, number of fuselage doors, weapons stations, and so on) are considered, it is apparent that the Black Horse airframe is potentially much lighter in weight than the F-16. The structural assumption pervading this analysis was that a 30,000 pound fighter aircraft at a normal load factor of 9 is similar, structurally, to a 180,000 pound aircraft of about the same size at a normal load factor of 1.5.
In the 1980s, Boeing Defense and Space Group developed an all-metal spaceplane concept called Reusable AeroSpace Vehicle, or RASV 11 . The aircraft was a large all-metal design that was subject to a large amount of detailed structural analysis later verified by structural demonstration. The Black Horse design has similar structural figures of merit to the RASV design, as seen in Table 3.
Table 3: Comparison of Black Horse and RASV
Parameter RASV Black Horse Wing area 5632 780 ft2 Wetted Area 19815 2194 ft2 Wing thickness 11.5 12 % Wing loading 217 237 lb/ft2 Wing Aspect Ratio 2.06 2.1 -- Swet/Sref 3.48 2.71 -- Subsonic L/D (max) 10 11 -- Engine wt/Dry wt 10 13 % Engine T/W 76 60 -- Dry wt/wetted area 7.9 7.64 lb/ft2 Dry wt/propellant volume 3.06 7.13 lb/ft3 Propellant mass fraction 87 91 % (TPS+Structure)/wet area 4.48 4.50 lb/ft2 Landing gear/landing wt 2.2 2.4 % Thrust/wt 0.726 0.767 --
The current Black Horse structural design needs much more detailed work. In particular, it would be advantageous to apply modern materials and design concepts to achieve the lowest weight structure possible. Advanced composites, in particular offer potential for great weight savings. The structural design of the NASA HL-20 space plane, which made extensive use of advanced composites 12 , had a structural weight per unit area of 1.34 lb/ft2. Aluminum/lithium alloys used in single-stage-to-orbit structural designs yield 2.0 lb/ft2 of aerostructure weight 13 . Indeed, many single-stage-to-orbit vehicles are sized by minimum gauge considerations to a great degree, permitting smaller vehicles to be even more robust if the panel weight is fixed. Finally, the NASA Ames research center HAVOC finite element vehicle sizing code 14 appears to substantiate the estimates for primary structure weight used in the initial study.
The concerns about the aerodynamic quality of the vehicle at the moment of tanker disconnect are largely correct. It was assumed in the early stages of analysis of the inflight propellant transfer concept that the optimal point for separation from the tanker was as high and as fast as the tanker could go. Additionally, the fact that the aircraft was in horizontal flight suggested that vehicle thrust to weight ratios of less than one were acceptable. Both of these assumptions are false, as further analysis has revealed.
The Program to Optimize Simulated Trajectories (POST) 15 was used to vary the engine thrust, drag, and separation altitude for the Black Horse vehicle, in an effort to examine the sensitivity of the aircraft's Delta-V to orbit and mass delivered to orbit to these variables. The basic Black Horse aerodynamic model was used, with lift held constant and drag varied as required.
Figure 1 Mass in 100 x 50 nmi orbit at 28.5
In Figure 1, the injected mass into orbit is presented for a number of separation altitudes and a propellant mass of 171,000 pounds. The higher thrust levels show decreasing returns in injected mass, indicating that at some point the additional mass in orbit is consumed by the additional engine weight. The reduction in wing area available by separating at a lower altitude may also exceed the injected mass lost by separating at a lower altitude. Drag was varied from 0.75 to 2.0 times nominal, with only a 150 pound variation in injected mass.
Using the structural numbers from Table 3 and requiring the lift coefficient to be 1.0 or less at tanker separation, it is possible to derive the results shown in Figure 2.
Figure 2 Black Horse Payload Mass with Lift Coefficient Constrained to 1.0
When the lift coefficient is constrained, the required wing area needed to fly at 43,000 feet consumes all the payload mass. At 30,000 feet, the payload bay volume begins to cut into the available propellant because the overall scale of the vehicle is reduced. Separation at 35,000 feet appears to offer the best compromise. The Delta-V required to achieve orbit in this case is 26,480 ft/sec.
The vehicle was resized for a payload to orbit of 1,000 pounds, yielding a propellant mass of 135,000 pounds and an injected mass of 12,975 pounds. The thrust was held constant at 210,000 pounds, reducing the Delta-v further to 26,250 ft/sec. The propellant mass fraction in this case is 91.9 percent. The wing area has been reduced to 620 ft2. The reduced propellant load should make it easier for the tanker to achieve the required altitude, and the reduction in oxidizer transfer to less than 120,000 pounds should reduce the amount of time on the tanker boom.
The required weight multiple remains a difficulty. The lift coefficient change is manageable, but the change in mass moments of inertia from hookup to disconnect will complicate the flight control design process. A weight multiple of 4.0 has been demonstrated in the B-52 16 , which is a challenging aircraft to refuel, but the high aspect ratio wing and lack of ailerons make B-52 refueling a completely different problem. A simulation of an aircraft with the linear small perturbation equations of motion was performed, and the mass moments of inertia were varied over a range consistent with a 4.0 weight multiple. An hx minimization flight control design was used. A predicted level one flying qualities rating was obtained for all gross weights. More attention needs to be paid to this area.
Further testing in a wind tunnel and in flight would be necessary to validate completely the aerodynamic assumptions made here. However, it is apparent that reducing separation altitude is beneficial structurally, aerodynamically and operationally.
The performance of the engine is indicated graphically in Figure 3. The NASA standard computer code TDK was used to predict the physical loss mechanisms (finite rate chemistry, two-dimensional nozzle, and boundary layer effects). The uncertainty in such measures is very small. The remaining correction that must be applied to these estimates to determine the delivered performance of the engine is the injector efficiency. Historically, H2O2/JP-5 engines have delivered over 98% efficiency with what would now be considered primitive injector designs. 17 , 18 The injector for the Black Horse engine has a two zone pattern, delivering a mixture ratio of 14.0 on the wall and 7.5 in the core. The core efficiency of the injector is 99.5%, based on calculations from the DROPMIX 19 computer code and actual injectors. Including the mixture ratio bias at the wall reduces the overall injector efficiency to 98%. The effect of the wall mixture ratio bias is to reduce the peak heat flux from 71 Btu/in2/sec to 46 Btu/in2/sec.
Figure 3 Engine performance and loss mechanisms
The engine is capable of operating at an expansion ratio of 240 at 43,000 feet without flow separation because it has a chamber pressure of 3,000 psi. However, the maximum heat flux measured for 98% hydrogen peroxide as a coolant is 44 Btu/in2/sec 20 . Furthermore, the requirement to reduce the wall heat flux is what drives the design to using a 98% efficient injector instead of the 99.5% efficient core design for the entire engine. If the core injector design is applied through the whole engine, the required expansion ratio to achieve 335 seconds of delivered vacuum Isp is reduced to below 140, as seen in Figure 3. To fit within the same fuselage diameter and be free of flow separation at 35,000 feet, the required chamber pressure can be reduced to 1,741 psi, and the expansion ratio is set to 145.9, assuming an optimized Rao nozzle. This causes the heat flux at the engine throat to be reduced to 32.5 Btu/in2/sec, within the limit for hydrogen peroxide coolant.
The reduction in chamber pressure has several beneficial effects. In addition to better operability and maintainability for the engine due to lower chamber pressure, the heat fluxes at the throat are now low enough that Narloy-Z copper alloy is no longer required. Inconel 718, which has performed well in the past with peroxide coolant for long-duration testing, should be sufficient. The reduced chamber pressure also makes the use of single stage pumps realistic. Furthermore, revised weight estimates of the engine have been performed 21 , showing a total installed engine weight of 2,200 pounds for 210,000 pounds of total thrust, based on the resized vehicle.
The turbine inlet conditions are somewhat moderated by the reduced preheating of the peroxide coolant, but remain stressing. Lower heat transfer to the turbine blades due to a reduction in inlet temperature to 1,780 F and a 2,000 psi reduction in turbine inlet pressure should make their design more tractable, but more work is needed in this area.
The initial Black Horse design was never intended to be anything more than an experimental proof-of-concept aircraft with some residual military utility. The fastest growing segment of the satellite market is the 1,000 pound payload class, and new satellite technologies have dramatically increased the capabilities of even very small satellites, such as MSTI and Clementine. Certainly increased payloads can be achieved by using expendable upper stages that are released exoatmospherically. It has also been suggested that a second exoatmospheric propellant transfer be performed 1 , which would permit substantially larger payloads or Delta-V. However, if weight growth or mission requirements drive the designer to making a larger aircraft or needing more Delta-V, it is worthwhile to ask how the concept scales to larger size. Also, alternative propellants, such as LO2/RP-1 or LO2/LH2 should also be examined. These require rejecting the KC-135 as the tanker of choice.
The first case examined was subject to the same assumptions as the Black Horse aircraft. Integral propellant tankage was assumed throughout, with 92, 90, and 88 percent of the volume to the outer skin of the aircraft available for propellant for H2O2/JP-5, O2/RP-1, and O2/H2 respectively. The payload mass was fixed at 1000 pounds, and the orbit was a 100 nmi circular orbit at 98 inclination. The vehicle performance is indicated in Table 4.
Table 4: Comparison of inflight propellant transfer spaceplane
weights for several propellants (1000 pounds payload) H2O2/JP O2/RP1 O2/H2 Propellant Mass 214174 239950 133428 lb Oxidizer Mass 188182 173900 112899 lb Empty Mass 17224 21776 22892 lb Structure 5837 7381 9976 lb Thermal Prot 2919 3691 4988 lb Landing Gear 1365 2381 1162 lb Takeoff Gross 62054 108217 52830 lb Overall Length 57.5 66.5 77.4 ft
Structural Mass at 2.6 lb/ft2
Thermal Protection System Mass at 1.3 lb/ft2
T/W at tanker release = 1.42
Miscellaneous Mass 2170 lb
Delta-V = 28,000 ft/sec, 35,000 ft separation altitude
100 x 100 nmi @ 98 orbit
Landing gear = 2.2% times takeoff gross weight
The preferred concept in terms of empty weight is the H2O2/JP-5 aircraft. This should probably result in the lowest development cost , especially since the aircraft has no cryogenic overhead. Notice, though, that the smallest amount of transfer propellant occurs for the largest aircraft.
A larger aircraft with a 10,000 pound payload requirement was also examined. The structural weights were raised somewhat, and the reference orbit was kept at 100 nmi circular at 98 inclination. The payload into a 28.5 inclination orbit was also calculated. The results are in table 5.
Table 5 Comparison of inflight propellant transfer spaceplane
weights for several propellants (10,000 pounds payload) H2O2/JP O2/RP1 O2/H2 Propellant Mass 629019 656029 344726 lb Oxidizer Mass 552680 476033 291693 lb Empty Mass 50587 59610 59144 lb Structure 14690 17584 22966 lb Thermal Prot 6078 7144 9330 lb Landing Gear 4010 6517 3003 lb Takeoff Gross 61591 296231 136494 lb Due east P/L 17773 19041 17532 lb Overall Length 85.3 92.6 105.8 ft
Structural Mass at 3.0 lb/ft2
Thermal Protection System Mass at 1.5 lb/ft2
T/W at tanker release = 1.42
Miscellaneous Mass 5000 lb
Delta-V = 28,000 ft/sec, 35,000 ft separation altitude
100 W 100 nmi @ 98 orbit
Landing gear = 2.2% times takeoff gross weight
The empty weights are fairly close together for this case, differing by only 15% among the various propellants. The preferred choice for this case is probably O2/H2. The wing loading is lowest, the landing gear weights are the lowest, and the transfer propellant requirement fits within the capacity of the KC-10 22 , rather than requiring a hypothetical new tanker. The difficulties of handling hydrogen and oxygen are probably worth the trouble for a vehicle with this level of capacity and performance. However, development costs would be much higher than those of the basic Black Horse experimental aircraft.
The HL-20 trajectory was used as a reference point only for the initial work. The actual reentry environment for the Black Horse aircraft should be much less stressing. Consider that the HL-20 design had a wing loading 23 at reentry of 66.6 lb/ft2, while the Black Horse design had a reentry wing loading of 20.9 lb/ft2. The Space Shuttle, by contrast, has a wing loading of over 100 lb/ft2. In an effort to examine the effects of the low wing loading design on reentry, the NASA Ames Research Center HAVOC code was used to perform a simulation. It was discovered that the peak temperature on the body during reentry was less than 2500 F, less than the 3080 F estimated for the HL-20 24 and observed on the Space Shuttle.
A detailed thermal protection sizing effort is needed to quantify the benefits of low wing loading reentry. Of particular interest is the leeside heating. It appears that about 70% of the upper surface of the Black Horse aircraft is exposed to temperatures of less than 700 F. The latest high temperature composite resins, such as AFR 700B 25 are certified for 100 hours of operation at such temperatures, are used in oxygenated environments such as engine exhausts, and are probably mature enough for use. Applying such materials offers the possibility of avoiding the thermal protection system altogether on up to 35% of the vehicle's surface.
consumption during transfer
It is tremendously inefficient to fly in the atmosphere for any length of time with a rocket engine, because the propellant consumption is about seven times that of an afterburning turbojet of the same thrust. As the weight of the aircraft increases, the required lift increases. The drag produced by this lift and hence the required thrust is a parabolic function of the aircraft's weight. The effect of this is to increase the consumption of oxidizer required just to maintain station on the tanker to a significant fraction of the transfer rate.
The propellant consumption was modeled for the resized Black Horse vehicle discussed in the previous section. The drag polar developed in the initial study was used to determine the required thrust as a function of aircraft weight. The fuel required for climb was included in the gross weight at tanker hookup, as was an allowance for additional fuel needed to burn with the oxidizer consumed during the transfer. The KC-10's transfer rate of 1500 gal/min was assumed, which would require changing the transfer pumps in the KC-135 (This would be required in any case for materials compatibility with H2O2). The aircraft is consuming propellant at almost 33% of transfer rate at the moment of disconnect. However, integrating the curves below shows that 82% of the transferred propellant is still on the receiver aircraft at the end of transfer, with the remaining 18% consumed. The total amount of propellant transferred is about 145,000 pounds, which falls within the KC-135's limits.
Figure 4 Propellant consumption during transfer (1500 gal/min)
Increasing the transfer rate would have a beneficial effect on reducing the oxidizer wasted during propellant transfer. The use of a larger diameter boom or a more effective transfer pump could reduce the time on the tanker below the 8 minutes in Figure 4, further improving the performance of the system and decreasing propellant waste. The information presented in Figure 4 was used to develop a mass history for the total mission. The assumptions were that climb from brake release to tanker altitude took two minutes of full-throttle operation, that rendezvous consumed 3 minutes at the best lift to drag ratio available at that altitude, and that the limit on axial load factor was 3.0. The results in Figure 5 show that the mission from brake release to orbital insertion takes about 20 minutes. This permits the aircraft to fly over any spot in the world within about an hour of brake release.
Figure 5 Black Horse Weight History
High-concentration hydrogen peroxide availability and
Hydrogen peroxide is not currently available at 98% concentration in the United States. It may, however, be produced from 70% concentration peroxide, which is a commodity item, by fractional crystallization 26 . The 70% liquid is chilled to -67 F, which forms a two-phase system consisting of solid hydrogen peroxide and 62% concentration liquid. The solid peroxide occludes a great deal of liquid, so centrifugal separation is required to yield the pure peroxide solid, which is then thawed. The process has several advantages. It is safer than distillation. Impurities tend to remain in the liquid solution rather than the solid precipitate. The 62% liquid may be distilled to 70% for reuse in the system. Most of the concern about 98% H2O2 for rocket applications is anecdotal. John Clark's book "Ignition -- An Informal History of Liquid Rocket Propulsion" devotes an entire chapter to hydrogen peroxide 27 , and has two problems with it. First, the freezing point is high -- about 31.4F (-0.4C). Anything added to hydrogen peroxide to depress the freezing point made it unstable and potentially explosive. Secondly, the Navy tested a puddle of jet fuel upon which they poured 90% H2O2. The peroxide sank though the fuel, began to decompose in contact with the dirt, formed an oxygen/fuel vapor mixture, and blew up. Spills of this type must, as a result, be avoided.
The second source of concern with peroxide is the Me-163B experience 28 and the capture of stocks of 70% H2O2 by the Allies after V-E day. The Me-163B often landed in flames and had a real problem with safety. The oxidant for the Me-163B was 70% H2O2, but it was manufactured by coerced labor with shoddy quality control under wartime conditions. Modern hydrogen peroxide, according to David Andrews 29 is a "completely different material". The Me-163B itself had wooden primary structure. Finally, the real risk was in the fuel -- a mixture of nitrous oxide, hydrazine hydrate, methanol, and potassium cuprocyanate. The Me-163A, which used 70% H2O2 as a monopropellant, was much safer.
The overwhelming choice for oxidizer in the aircraft rocket world has been hydrogen peroxide. The AR-2 engine, used in the FJ-4, F-86, and the NF-104, was a 90% H2O2 and JP-5 or JP-4 engine, had a two hour time between overhauls (a number that isn't even specified for most rockets) and was operated and maintained by ordinary Air Force enlisted servicemen for years 18 . The Snarler and Screamer engines used in the UK's Buccaneer fighter also employed 85% H2O2 and kerosene 17 , and eventually begat the Gamma engine used in the Black Knight and Black Arrow rocket programs.
Hydrogen peroxide in any concentration is an oxidant and as such needs to be treated with respect and care. It is clearly a less powerful oxidant than oxygen, but even so, it has to be handled according to a well-defined set of procedures. The hazards usually manifest themselves in the effect of impurities on the peroxide rather than the effect of the peroxide on the impurities. Notice that this is the reverse of the mechanism of failure with liquid oxygen, where a small impurity tends to burn and cause an evolution of oxygen gas that destroys delicate parts and leads to catastrophic failure. Nevertheless, the failures are equally catastrophic and the standard of cleanliness is the same. Impurities of all kinds, particularly organics, must be absolutely avoided 30 .
One additional precaution is needed with peroxide -- anything that touches it must be passivated beforehand. There are a large number of procedures for passivation, generally involving the washing of the part with high strength nitric acid and then with progressively higher grades of peroxide until final peroxide strength is reached. Not all materials are suitable for peroxide use. Stainless steels, some aluminum alloys, zirconium, glass, and tin can all be treated to class 1 compatibility with 98% H2O2. Class 1 means "suitable for storage tanks and long term continuous exposure" and involves a decomposition rate of 0.4 to 0.1 % per year. Of particular concern is the choice of materials for lubricants and seals. Only fluorinated polymers (such as Teflon, Kel-F, or Viton) appear to be suitable.
An interesting result of the long term compatibility results for hydrogen peroxide is that 98% H2O2 is more stable that 90% H2O2. The reason for this appears to be that the water molecule is very slightly catalytic, being polar 31 . Also, elevated pressure can suppress decomposition (by reason of Le Chatelier's principle), but the recommended practice is to vent peroxide storage and transport containers.
The hazards of dealing with high purity hydrogen peroxide fall into four categories: detonation and explosion, uncontrolled decomposition, fire, and personnel injury.
Concentrated vapors will irritate the nasal passages and eyes. Vapors, mists, and liquid will irritate skin. Ingested peroxide will decompose internally, leading to severe distention of the stomach and internal burns. The corrective action is to flush with water. Do not ingest. The vapor pressure is only 1/9 that of water, which helps prevent harmful exposure to hazardous vapor levels.
Hydrogen peroxide is not flammable, but will react with combustible materials with the evolution of enough heat to initiate and support combustion. Removing the air does not help, because the peroxide generates its own oxygen on decomposition. Fight peroxide fires with water. Chemical extinguishers will catalyze further decomposition.
98% H2O2 is not classified as impact or shock sensitive. Numerous tests have been unable to sustain detonation waves in liquid peroxide solutions. Vapor phase concentrations of over 26% peroxide are considered "explosive" in the sense that the release of energy in the vapor phase upon decomposition is rapid enough to produce effects normally associated with explosions. For 98% H2O2 the limit 32 is 212 F. Invariably, peroxide vapor hazards are preceded by a slow buildup of temperature and pressure in the tanks. The corrective action is to monitor temperature and pressure buildups, vent the tanks, do not permit elevated temperatures and avoid impurities.
98% H2O2 can be, and indeed has been, a safe and effective rocket propellant PROVIDED THE RIGHT DESIGN, MANUFACTURE, AND OPERATIONS PROCEDURES ARE FOLLOWED . The entire system must be composed of peroxide compatible materials, preferably class 1. The system must be designed and operated in such a way as to prevent contamination with reactive materials (no garden hose purges, no greasy handprints on the refueling nozzle, etc.). Keep the number of mechanical joints to a minimum. Vent ball valves upstream. Avoid threaded connections. Design the system to amply sustain the maximum operating pressure. Avoid liquid traps in propellant lines. The purge system must not require disconnecting any system joints. All components must be reliable, compatible with peroxide, and properly cleaned and passivated. Following these procedures can assure the user of first-time safety and success 33 .
Using hydrogen peroxide as a rocket oxidant can offer significant benefits provided it can be handled and used safely. This is a paramount issue for modern rocket designers, and the exothermic nature of 98% H2O2 causes some legitimate concerns. The solution to these concerns is not in new technology, but in proper design, manufacture and operations procedures -- in short, the answer is discipline.
The inflight propellant transfer concept offers a great degree of capability and flexibility, with little in the way of required technological development. It can be used with a variety of different propellant combinations, the selection of which depends on scale and other mission requirements. The concept is arguably easier to test than competing concepts and uses many existing resources, such as tankers and runways, that have already been paid for.
Many criticisms have been made of the initial Black Horse design, as a result of which the concept has become stronger and more credible.