The following presentation was given at a NASA sponsored conference at the University of Alabama in Huntsville during the week of February 23 1998. The presentation shown here was cleared for public release and was included in the published proceedings for that conference.
Five years ago the Skunk Works began a long term project whose purpose is to explore the feasibility of designing, building and operating a market driven, commercially funded, commercial launch service that is affordable to build, profitable to operate, low enough in user cost that it allows the market to grow, and can grow with the market.
The only technology constraint placed on the program is that it utilize existing or near term technology. There are no requirements regarding propulsion systems, propellant types, operating modes, or number of stages placed on this study. It is our thought that the requirements of minimizing the size of the initial investment, operating at a profit with the lowest possible user cost, and having built-in evolutionary growth potential would be more than adequate as selection criteria.
In Phase 1 of this study we investigated rocket-powered launch vehicles. Included in this were Vertical Take-Off Horizontal Landing (VTOHL), Vertical Take-Off Vertical Landing (VTOVL), and Horizontal Take-Off Horizontal Landing (HTOHL). All three were examined for both Single Stage To Orbit (SSTO) and Two Stage To Orbit (TSTO) flight. The results of this work were turned over to what became the X-33 program.
Since then we have been investigating air-breathing launch vehicle. As before, we will be investigating VTOHL, VTOVL, and HTOHL. All three will be examined for SSTO and TSTO flight.
The Next Generation Reusable Launch Vehicle
We started with the HTOHL concept shown here. It is what we call a stage and a half concept since it has a small expendable upper stage. We selected this concept because all of the HTO concepts designed for SSTO flight were too large for existing commercial facilities -- a problem which tends to defeat the purpose of selecting a Horizontal Take-Off vehicle in the first place.
We also found that the HTO SSTO vehicles were significantly more expensive to build and operate then the stage and a half concept. This was due to the larger size and correspondingly larger initial investment needed for the SSTO vehicle, the low flight rate required by the existing launch market, and the fact that we included debt service in the cost model.
User Cost as a Function of Staging Velocity
This chart shows exactly how large an impact, changes in the staging velocity can have on the user cost. This particular chart, which was developed in Phase 1 of this study, is for an unmanned prototype, Vertical Take-Off Horizontal Landing (VTOHL) lifting body, that has two Space Shuttle Main Engines (SSME) rocket engines and a Gross Take-Off Weight (GTOW) of 600,000 pounds. User cost is shown on the Y-axis and useful payload to Low Earth Orbit (LEO) on the X-axis. Staging velocities are shown as a percentage of orbital velocity at each data point on the curve.
As shown on the chart, user cost goes down dramatically as the staging velocity goes down. This is due to the increasing size of the useful payload, which provides a larger base to distribute the launch costs over. At a staging velocity of around 65% to 70% the user cost starts to go back up. This is due to the increasing cost of the expendable upper stage, which begins to overpower the cost savings generated by the increasing payload mass at these staging velocities.
This relationship between user cost and staging velocity also applies to the air breathing launch vehicles since the phenomena is driven more by the inclusion of debt service in the model then any hardware differences.
The upper curve shows the user cost when all private money is used to develop the launch vehicle. The lower curve shows what happens to the user cost when $600 million dollars of the initial investment comes from the government. We refer to government money in this study as ZPI, or Zero Payback Investment. We assume for this study that ZPI does not have to be paid back - hence the term Zero Payback Investment.
Note: Before comparing this vehicle with the air breathing concept being presented, it is necessary to keep in mind that these cost estimates do not include the cost of ferrying the reusable first stage back to its launch site, nor do they allow for the real world operational restrictions on staging velocity that will occur when selecting a landing site for the reusable first stage for a given flight path. Both of these problems will add hundreds of dollars per pound to the user cost.
Also, this rocket powered vehicle concept is Vertical Take-Off (VTO) while the air breather is Horizontal Take-Off (HTO). A more valid comparison would be with a VTO air breather or a HTO rocket. The fact that the HTO air breather is user cost competitive with a VTO rocket would seem to indicate that a VTO air breather would be noticeably better then a VTO rocket.
ZPI and Flight Rate
Something else we discovered about ZPI was its favorable impact on user cost as a function of flight rate. As shown in the chart, the upper curve shows user cost as a function of flight rate when all private money is used for the initial investment. This cost curve puts the initial users in a higher cost bracket then those who sign on later when the risk of flying on a new launch vehicle is less. The lower curve shows the impact of increasing flight rate on user cost when the optimum amount of government money is included in the initial investment.
Payload Size and Flight Rate
During Phase 1 of this study when we were investigating rocket-powered reusable launch vehicles, a market study was performed to determine the optimum payload size and flight rate that maximized market share and minimized user cost. Two of the prime factors that contribute to a minimum user cost are minimizing the size of the initial investment and maximizing the flight rate. This drove us towards the smallest possible launch vehicle that could carry the largest number of payloads.
In order to try and get a better understanding of this, an investigation was performed to determine the number of payloads as a function of their delivered mass to Low Earth Orbit (LEO). This investigation found that 80% of all commercial payloads could be carried by a launch vehicle with a payload capacity of 15,000 to 16,000 pounds.
A more recent preliminary design study of a 6-person Personnel Launch Vehicle for crew exchange with the International Space Station led us to increase that payload size to 17,500 pounds.
Since the current total commercial launch market consists of approximately 800,000 pounds a year to LEO, 80% of that comes out to approximately 36 flights per year at a payload size of 17,500 pounds. A flight rate that could be served by 2 launch vehicles flying 18 flights per year each.
Personnel Launch Vehicle, Ascent
One of the questions people have had regarding this launch vehicle is how it can be used for crew exchange with the International Space Station. This chart shows the ascent flight profile for a six-man Personnel Launch Vehicle (PLV) that is designed for that mission.
The launch vehicle carries the PLV with attached expendable upper stage to an altitude of 150 km and a velocity of approximately 70% of orbital velocity for that altitude. Shortly after separation from the launch vehicle, the expendable upper stage is ignited in order to accelerate the PLV the rest of the way to orbital velocity. The expendable upper stage is then jettisoned, de-orbited, and allowed to burn up in the Earth's atmosphere. Upon achieving the correct orbital alignment with the International Space Station, the PLV then uses its onboard propulsion system to complete the rendezvous.
Personnel Launch Vehicle, Descent
For the return flight, the PLV undocks from the International Space Station and de-orbits using its on-board propulsion system. After re-entry, and once it has slowed to subsonic speeds, the parachutes are deployed and the landing gear struts extended. Just before touch-down the PLV then fires its landing rockets in order to further reduce its landing velocity to less then 10 feet per second.
Propulsion System Studies
Our number one goal at this time is to map out the size and cost impacts of the various air breathing propulsion system combinations on the HTOHL vehicle. So far we have identified 3 basic propulsion system categories with at least 6 sub-categories in each.
Those three basic categories are;
Turbo-Rocket Combined Cycle, or TRCC,
Turbine Based Combined Cycle, or TBCC, and
Rocket Based Combined Cycle, or RBCC.
Turbo-Rocket Combined Cycle (TRCC)
This chart shows the basic arrangement of the Turbo-Rocket Combined Cycle engine concept. It consists of a Ramjet-Scramjet, multiple ducted rocket motors, and multiple parallel flow path afterburning turbofan engines.
TRCC
This chart shows the range of possible engine operating modes that are possible with this type of propulsion system.
As shown, this vehicle takes-off and accelerates to approximately Mach 2.5 using the parallel flow path afterburning turbofan engines and the liquid oxygen and hydrocarbon ducted rocket motors that are inside the ramjet-scramjet flow path. Additional thrust is generated by the ramjet starting at approximately Mach 0.8.
At Mach 2.5 the turbofans are shutdown and the vehicle continues to accelerate using only the liquid hydrogen fueled ramjet-scramjet.
Somewhere between Mach 8 and Mach 10 the ducted rocket motors are re-ignited with liquid oxygen and liquid hydrogen. The vehicle then continues to accelerate using both the scramjet and the ducted rocket to somewhere between Mach 14 and Mach 18 when both are shutdown. The vehicle will then coast up to its staging altitude of 150 km where the payload with expendable upper stage is released.