In-flight propellant transfer spaceplane design and testing
considerations
In-flight propellant transfer spaceplane design and testing
considerations
Captain Mitchell Burnside Clapp
Phillips Laboratory, Kirtland AFB, NM 87117
Abstract
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 an extremely 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. 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 400 manned aircraft have ever been released
from beneath other aircraft. 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.
Black Horse
In order to examine the utility of the inflight propellant
transfer concept, and its application to military requirements, a
contracted six-week study 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:
Horizontal takeoff like an aircraft
Two engines firing at takeoff
Propellant transfer as high and as fast as possible within the
envelope of the KC-135Q tanker
Hydrogen peroxide and jet fuel propellants
Power-off landing
LEO mission (100 x 50 nmi @ 28.5 inclination)
Throttling during propellant transfer
Maximize use of existing facilities and support equipment
Conservative design assumptions
Propellant Selection
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 spe-
cial 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 of a hydrogen peroxide/jet fuel engine indi-
cates 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 (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
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.
Mission Profile
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.
Weights
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 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 con-
ventional structural technology by forming the blended wing/body
of the aircraft from ordinary aluminum alloy. The thermal protec-
tion 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.
Design Considerations
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 take-
off or landing, especially with the enormous thrust available
from the rocket engine. Low wing loading may also moderate the
thermal environment during reentry.
Flight test
Unlike most space vehicles, it will be possible to test the air-
craft 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 7.30:1, 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 4800 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 Delta-V 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.
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.
Structural weight
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.
Aerodynamics
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.
Engine performance
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.
Size concerns
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.
Aerothermodynamics
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
active cooling of leading edges, and with no cryogens aboard the
aircraft it is not clear how this can be accomplished.
Propellant 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 make 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
Structural weight
The structural weight estimates were developed using statistical
weight estimation methodology. 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. 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 below 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, 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. 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 appears to substantiate the estimates for primary structure
weight used in the initial study.
Aerodynamics
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) 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 propellant mass fraction in this case is
91.8 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 mass-to-altitude performance, 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, 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. 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.
Engine performance
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. 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 computer code and comparisons with actual injectors
with the same pattern. 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 long duration (> 10
minutes) heat flux measured for 98% hydrogen peroxide as a
coolant is 34 Btu/in2/sec. 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 150, 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. This causes the heat flux at the engine
throat to be reduced to 32.5 Btu/in2/sec, well 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.
Furthermore, revised weight estimates of the engine have been
performed, 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
pressure should make their design more tractable, but more work
is needed in this area.
Size concerns
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, 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
Assumptions:
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
Assumptions:
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, 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.
Aerothermodynamics
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 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 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 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.
Propellant 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.
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 handling
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. The 70% liquid is chilled to -67 F, which forms
a two-phase system consisting of solid hydrogen peroxide and 62%
concentration liquid. The slush of solid peroxide occludes a
great deal of liquid, so centrifugal separation is required to
yield the peroxide solid, which is then thawed and stored. 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, or sold back to the initial peroxide vendor.
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, 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 strictly avoided.
The second source of concern with peroxide is the Me-163B
experience 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 ("Advantages of Hydrogen Peroxide as a
Rocket Oxidant", JBIS, Jul 1990) 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
enormously 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. The Snarler and
Screamer engines used in the UK's Buccaneer fighter also employed
85% H2O2 and kerosene, 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.
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 % concentration loss per year. Of particular concern is the
choice of materials for lubricants and seals. Only fluorinated
polymers (Teflon, Kel-F, Aclar, and 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 (Roth, E. M., and Shanley, E. S.,
"Stability of pure hydrogen peroxide, " Ind. Eng. Chem, 45, 2343-
9, 1953). 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 the following 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 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.
The application of 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.
Conclusions
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.