A LO2/Kerosene SSTO Rocket Design Study

In February of 1997, Mitchell Burnside-Clapp posted a vehicle dsign study to the sci.space.policy Usenet Newsgroup. It was based on a NASA Access To Space study and looked at the same vehicle design but using higher density LO2/Kerosene propellants and engines rather than the LOX/Liuid Hydrogen propellant and engines of NASA's design. The results were quite interesting.

A LO2/Kerosene SSTO Rocket Design (long)

Mitchell Burnside Clapp
Pioneer Rocketplane

(view with a fixed pitch font such as courier or monaco)


The NASA Access to Space LO2/hydrogen single stage to orbit rocket was examined, and the configuration reaccomplished
with LO2/kerosene as the propellants. Four major changes were made in assumptions. First, the aerodynamic
configuration was changed from a wing with winglets to a swept wing with vertical tail. The delta-V for ascent
was as a result recalculated, yielding a lower value due to different values for drag and gravity losses. The engines
were changed to LO2/kerosene burning NK-33 engines, which have a much lower Isp than SSME-type engines used in the
access to space study, but also have a much higher thrust-to-weight ratio. The orbital maneuvering system on
the Access to Space Vehicle was replaced with a pump-fed system based on the D-58 engine used for that purpose now
on Proton stage 4 and Buran. Finally, the wing of the vehicle was allowed to be wet with fuel, which is a
reasonable practice with kerosene but more controversial with oxygen or hydrogen. Additionally, in order to reduce
the technology development needed, the unit weights of the tankage were allowed to increase by 17 percent.

After the design was closed and all the weights recalculated, the empty weight of the LO2/kerosene vehicle
was 35.6% lighter than its hydrogen fuelled counterpart.


NASA completed a study in 1993 called Access to Space, the purpose of which was to consider what sort of vehicle
should be operated to meet civil space needs in the future.  The study had three teams to evaluate three different broad
categories of options. The Option 3 team eventually settled on a configuration called the SSTO/R. This vehicle was a
LO2/hydrogen vertical takeoff horizontal landing rocket.  The mission of the Access to Space vehicle was to place a
25,000 pound payload in a 220 n.mi. orbit inclined at 51.6 degrees. The vehicle had a gross liftoff weight of about
2.35 million pounds. The thrust at liftoff was 2.95 million pounds, for a takeoff thrust to weight ratio of 1.2. The
empty weight of the vehicle was 222,582 pounds, and the propellant mass fraction (defined here as [GLOW-empty]/GLOW) 
was 90.5%.

Main power for this vehicle was provided by seven SSME derivative engines, with the nozzle expansion ratio reduced
to 50. This resulted in an Isp reduction from 454 to 447.3 seconds. Each engine weighed 6,790 lbs, for an engine sea
level thrust to weight ratio of 62.

Aerodynamically the vehicle was fairly squat, with a fineness ratio (length:diameter) of 5. The overall length
of the vehicle was 173 feet and its diameter was 34.6 feet.  It had a single main wing (dry of all propellants) of about
4,200 square feet total area, augmented by winglets for directional control at reentry. The landing wing loading
was about 60 lb/ft2. The oxygen tank was in the nose section. The payload was mounted transversely between the
oxygen and hydrogen tanks, and was 15 feet in diameter and 30 feet long.

This design exercise was among the most thorough ever conducted of a single stage to orbit LO2/LH2 VTHL rocket.
It was probably the single greatest factor in convincing the space agency that single stage to orbit flight was
feasible and practical, to borrow from the title of Ivan Bekey's paper of the same name.

A LO2/kerosene alternative

A number of people have been asserting for some time that higher propellant mass fractions available from dense
propellants may make single stage to orbit possible with those propellants also. The historical examples of the
extraordinary mass fractions of the Titan II first stage, the Atlas, and the Saturn first stage are all persuasive.
Further, denser propellants lead to higher engine thrust to weight ratios, for perfectly understandable hydraulic

It has not usually been observed that higher density also leads to significant reductions in required delta-v.
There are two major reasons that this is so. First, the reduction in volume leads to a smaller frontal area and
lower drag losses. The second, and more significant, reason is that the gravity losses are also reduced. This is because
the mass of the vehicle declines more rapidly from its initial value. The gravity losses are proportional to the
mass of the vehicle at any given time, and hence the vehicle reaches its limit acceleration speed faster.

NASA itself has implicitly recognized this effect. When the Access to Space Option 3 team examined tripropellant
vehicles, the delta-v to orbit derived from their work was 29,127 ft/sec, for precisely the reasons described in the
previous paragraph. This compares to a delta-v of 30,146 ft/s for the hydrogen-only baseline, as reported in a
briefing by David Anderson of NASA MSFC dated 6 October 1993. To be clear, these delta-v numbers include the back
pressure losses, so that no "trajectory averaged Isp" number is used. They did not, however, report any results
for kerosene-only configurations.

To come to a more thorough understanding of the issues involved in SSTO design, I have used the same methodology
as the Access to Space team to develop compatible numbers for a LO2/kerosene SSTO. There are four major changes in
basic assumption between the two approaches, which I will identify and justify here:

1: The ascent delta-v for the LO2/kerosene vehicle is 29,100 ft/sec, rather than 29,970 ft/sec. The reason for
this is argued above, but I ran POST to verify this value, just to be sure. The target orbit is the same: 220 n.mi.
circular at 51.6 degrees inclination. The detailed weights I have for the NASA vehicle are based on a delta-v of
29,970 ft/sec rather than the 30,146 ft/sec reported in Anderson's work, but I prefer to use the values more
favourable to the hydrogen case to be conservative. The optimum value of thrust to weight ratio turns out to be
slightly less than the hydrogen vehicle: 1.15 instead of 1.20.

2: The aerodynamic configuration is that of Boeing's RASV.  Without arguing whether this is optimal, the fineness ratio
of 8.27 and large wing lead to a much more airplane-like layout, better glide and crossrange performance, and
reduced risk. The single vertical tail is simpler and safer than winglets as well. Extensive analysis has justified the
reentry characterisitics of this aircraft. The wing is assumed to be wet with the kerosene fuel, as is common on
most aircraft. The fuel is also present in the wing carry-through box. The payload is carried over the wing
box, and the oxidizer tank is over the wing. This avoids the need for an intertank, which in the NASA Access to
Space design is nearly 6,600 pounds.

3. The main propulsion system is the NK-33. The engine has a sea level thrust of 339,416 lbs, a weight of 2,725 lbs
with gimbal, and a vacuum Isp of 331 seconds. Furthermore, it requires a kerosene inlet pressure of only 2 psi
absolute, which dramatically reduces the pressure required in the wing tank. It also operates with a LO2 pressure at
the inlet of only 32 psi. The comparable values for the SSME are about 50 psi for both propellants. This will have
a substantial effect on the pressurization system weight.

4. The OMS weight is based on the D-58 engine. This engine is used for the Buran OMS system and the Proton stage 4. As
heavy as it is the Isp is an impressive 354 seconds. NASA's vehicle used a pressure fed OMS, which is a sensible design
choice if you're stuck with hydrogen and you wish to minimize the number of fluids aboard the vehicle. But
because both oxygen and kerosene are space-storable, there is no reason to burden the design with a heavy pressure fed

Using the same methodology for calculating masses, and accepting the subsystems masses as given in the Access to
Space vehicle, a redesign with oxygen and kerosene was accomplished. The results appear in Table 1.

Table 1: Access to Space vehicle and LO2/kerosene
Name O2/H2 LO2/RP Wing 11,465 11,893 lb Tail 1,577 1,636 lb Body 64,748 33,741 lb Fuel tank 30,668 - lb Oxygen tank 13,273 17,271 lb Basic Structure 14,610 10,274 lb Secondary Structure 6,197 6,197 lb Thermal Protection 31,098 21,238 lb Undercarriage, aux. sys 7,548 5,097 lb Propulsion, Main 63,634 36,426 lb Propulsion, RCS 3,627 1,234 lb Propulsion, OMS 2,280 823 lb Prime Power 2,339 2,339 lb Power conversion & dist. 5,830 5,830 lb Control Surface Actuation 1,549 1,549 lb Avionics 1,314 1,314 lb Environmental Control 2,457 2,457 lb Margin 23,116 16,105 lb Empty Weight 222,582 141,682 lb Payload 25,000 25,000 lb Residual Fluids 2,264 1,911 lb OMS and RCS 1,614 1,261 lb Subsystems 650 650 lb Reserves 7,215 8,895 lb Ascent 5,699 7,587 lb OMS 679 541 lb RCS 837 767 lb Inflight losses 13,254 17,445 lb Ascent Residuals 10,984 15,175 lb Fuel Cell Reactants 1,612 1,612 lb Evaporator water supply 658 658 lb Propellant, main 2,054,612 3,034,972 lb Fuel 293,604 843,048 lb Oxygen 1,761,008 2,191,924 lb Propellant, RCS 2,814 2,556 lb Orbital 2,051 1,756 lb Entry 763 800 lb Propellant, OMS 19,357 15,452 lb GLOW 2,347,098 3,246,156 lb Inserted Weight 292,486 211,185 lb Pre-OMS weight 271,482 186,152 lb Pre-entry Weight 252,125 170,700 lb Landed Weight 251,362 169,900 lb Empty weight 222,582 141,682 lb Sea Level Thrust 2,816,518 3,733,080 lb Percent margin 11.6% 12.8% Assumed Isp(vac) 447.3 331.0 s Ascent Delta-V 29,970 29,100 ft/s OMS delta-V 1,065 987 ft/s RCS delta-V 108 107 ft/s Deorbit Delta-V 44 53 ft/s Reserves 0.28% 0.25% lb/lb Residuals 0.53% 0.50% lb/lb Wing Parameter 4.56% 7.00% lb/lb TPS parameter 12.37% 12.50% lb/lb Undercarriage parameter 3.00% 3.00% lb/lb Wing Reference Area 4,189 5,528 ft2 Density of fuel 4.4 50.5 lb/ft3 Density of oxygen 71.2 71.2 lb/ft3 Volume of fuel 66,276 16,694 ft3 Volume of oxygen 24,733 30,785 ft3 Fuel tank parameter 0.42 - lb/ft3 Oxygen tank parameter 0.48 0.56 lb/ft3
Some discussion of the results and justification is in order. The wing is about 40 percent heavier as a percentage of landed weight than for the hydrogen fueled baseline. When considered as a tank, it is about 60 percent heavier for the volume of fuel it encloses. Its weight per exposed area is about the same and the wing loading is half at landing. No benefit is taken explicitly for the lack of a requirement for kerosene tank cryogenic insulation. The tail is assumed to have the same proportion of wing weight for both cases. This is conservative for the kerosene wehicle because its single vertical tail is structurally more efficient. The body of the kerosene vehicle has three components. The oxidizer tank has an increased unit weight of about 17 percent. This is done in order to avoid the need for aluminum-lithium, which was assumed in the Access to Space vehicle. The basic structure group is unchanged, except that the intertank is deleted and the thrust structure is increased in proportion to the change in thrust level. The secondary structure group is mostly payload support related, and was not changed. The thermal protection group is in both cases about 12.5% of the entry weight. This works out to 1.107 lbs/ft2 of wetted area for the kerosene vehicle, which is common to many SSTO designs. The undercarriage group is 3% of landed weight for both vehicles. There is no benefit taken for reductions in gear loads for the kerosene vehicle due to lower landing speed and lower glide angle at landing. The main propulsion group includes engines, base mounted heat shield, and pressurization/feed weights. The engines are far lighter for their thrust than SSME derivatives. The pressurization weights are reduced in proportion to the pressurized volume for the kerosene vehicle. No benefit is taken for reduced tank pressure. Here is as good a place as any to point out the erroneous assertion that increased hydrostatic pressure is going to lead to increased tankage weights. There is no requirement for a particular ullage pressure except for the need to keep the propellants liquid. It is the pressure at the base of the fluid column rather than the top of the column that is of engineering interest. The column of fluid exerts a hydrostatic load on the base of the tank, but this load does not typically exceed the much more adverse requirement for engine inlet pressurization. For the kerosene vehicle, the hydrostatic load at the base of the oxygen tank is 49 psi, which is compatible with the pressures normally seen in oxygen tanks for rocket use. The load declines after launch because the weight goes down faster than the acceleration goes up. The bottom line here is that dense propellants may require you to alter a tank's pressurization schedule, but not to overdesign the entire tank. Structures are sized by loads and tankage for rockets is sized principally by volume, and if the vehicle is small, by minimum gauge considerations. This is not completely true for wet wings, however, as discussed previously. In this particular example, there is no need for high pressure in the wing tank either, because of the low inlet pressure required by the NK-33. The OMS group is the only other major change, as discussed above. The reliable D-58 engine has been performing space starts for decades and will serve well here. The acceleration available from the OMS is about 0.12 g, which is standard. All the other weights are pushed straight across for the most part. A brief inspection suggests that this is very conservative. Control surface actuation requirements are certainly less, electrical power requirements less, much better fuel cells available than the phosporic acid type assumed here, and reduced need for environmental control. Nonetheless, rather than dispute any of these values it is easier simply to accept them. The margin is applied to all weight items at 15% execpt for the engine group at 7.5%. The justification for this is that the main and OMS engine weights are known to high accuracy. The vehicle has an overall length of 1955 inches, and a diameter of 236.4 inches. The wing has a leading edge sweep of 55.5 degrees and a trailing edge sweep of -4.5 degrees. Its reference area is 5,632 square feet, of which 3,992 square feet is exposed. The wing encloses 16,694 ft3 of fuel, with a further 5% ullage. The carry-through is also wet with fuel. The wing span is 1293 inches, and the taper ratio is 0.13. The payload bay has a maximum width and height of 15 feet. It sits on top of the wing carry through box. The thrust structure from the engines passes through and around the payload bay to the forward LO2 tank. The payload bay is 30 feet in length. It has a pair of doors, the aft edge of which is just forward of the vertical tail leading edge. The engine section encloses 11 NK-33 engines, with a 4 - 3 - 4 layout. The engines are each 12.5 feet long, and additional structure and subsystems take up another 6.5 feet. The oxygen tank comprises the forward fuselage, which encloses 30,785 ft3 of oxygen, with a further 5% ullage. The length of the tank is about 100 feet. The ventral surface of the tank is moderately flattened as it moves aft, to fair smoothly with the wing lower surface. This flattening reduces its length by about 5% with respect to a strictly cylindrical layout. The aft edge of the oxygen tank is about even with the forward payload bay bulkhead. A compartment of about 13.9 feet provides room for some subsystems and a potential cockpit in future versions. Conclusion The methods of the NASA Access to Space study were used to design a single stage to orbit vehicle using existing LO2/kerosene engines. An inspection of the final results shows that the vehicle weighs about 36.5% less than its hydrogen counterpart, with reductions in required technology level and off the shelf engines. The center of mass of the vehicle is about 61% of body length rather than 68% for the Access to Space vehicle, which should improve control during reentry. The landing safety is considerably improved by lower landing speed and better glide ratio. Structural margins are greater overall. The vehicle designed here appears to be superior in every respect: smaller, lighter, lower required technology, improved safety, and almost certainly lower development and operations cost.