This section will explore the short-term options for reaching orbit and the International Space Station (IIS), the options to replace the Space Shuttle by the end of the decade, and follow-on methods for reducing the cost of human access to space with the goal of building a new reusable space infrastructure.
Currently, the only alternative to the Space Shuttle is the Russian Soyuz capsule. While the Soyuz and Progress infrastructure will allow limited occupation of the ISS as currently configured, it will not allow for the completion of station construction, or for an expanded crew to support research opportunities. The only choice for the next five years is to work through the Columbia investigation, make the necessary technical and procedural changes and fly the remaining three shuttles.
The current ISS construction schedule requires 25-30 shuttle flights over 5-6 years to complete full station construction and support operations. The only non-IIS mission on the manifest is a Hubble servicing mission (STS-122). There are 23 listed IIS shuttle missions and a further few would be necessary for the addition of a habitat module and possible pre-positioning of supplies and equipment mandated by post-Columbia modifications. This translates into approximately 10 missions for each remaining shuttle in the 2004-2009 timeframe.
The Space Shuttle seems a flawed space transportation system. At best, it is not a cost-effective method of launching people and cargo to orbit. At worst, it is also unsafe and unreliable. While it was perhaps inevitable that the shuttle program would continue and Endeavour would join the fleet after the Challenger disaster on the 25th flight, a similar decision would be disastrous after the 113th flight. It is in NASA's best interest to replace the shuttle as quickly as possible. The fastest method to accomplish that goal is through the use of expendable launch vehicles for human and cargo launches.
For the purposes of the following discussion the Delta IV family is used as a baseline. This is not to slight the Atlas V or Ariane V vehicles, but because of better access to relevant performance data from public sources for the Delta vehicles.
NASA's stated direction is to develop an Orbital Space Plane (OSP) as an IIS lifeboat by 2010 and as a crew transfer vehicle by 2012. Accelerating this effort can produce a viable crew transfer vehicle by 2009 to completely replace the shuttle. Detailed below is the Aquila spaceplane, a proposed configuration exceeding the stated Level I requirements for the OSP.
The Aquila is loosely based on the NASA Langley HL-20 design. Massing 11 metric tonnes (mt), including maneuvering fuel and a solid rocket abort system, the Aquila could deliver a crew of six or more to the IIS atop a Delta IV Medium 4,2 launch vehicle or reach the Earth-Moon Lagrange 1 point (L1) atop a Delta IV Heavy vehicle. The figure below details component masses:
|Aquila Space Plane||mt||delta v||ISP||Notes|
|Actual Liftoff Mass||11.000|
|Airframe||3.500||Includes built-in tanks|
|Abort Engines (x4)||0.200||22,000 kgf each; expendable; usable for orbital circulation and de-orbit|
|Abort Fuel - 240m/s||0.940||
|Solid Fuel - 3 second full burn - 4 g per engine pair|
|OMS Engines (x2)||0.240||2,000 kgf thrust each - based on RL10 design|
|OMS Fuel - 900m/s||2.400||
|CH4/O2 Fuel for compact storage and common cooling|
|RCS Thrusters (x25)||0.100||100 N thrusters|
|RCS Fuel - 100m/s||0.350||
|Landing Gear||0.550||5% loaded mass -7.5% landing mass|
|Avionics||0.400||Controls, computers, communications|
|Rear Airlock Doors||0.500||Separately pressurized with docking and crew access doors|
|Crew Stations (x6)||0.600||Occupied weight including personal items|
|EVA Suits (x3)||0.195||Orlan based or EMU based - 65kg each|
|Fuel Cell Units (X2)||0.100||10kg/kw - 10kw max|
|Fuel Cell Fuel||0.600||60kg/day - 10 days full power - 20 at half power: produces onboard H2O|
|Margin/Cargo||0.325||consumables and spares|
The Aquila design is a lifting body 8.5m in length, with a 2m wide crew section seating at least six astronauts. To minimize total system mass, four expendable solid rocket abort engines, producing 22mt of thrust each for a three second burn provide abort capabilities (an 8g burn to clear the launch vehicle). These engines, used in pairs can, can provide the thrust for the initial orbital circulation burn and for the deorbit burn, saving maneuvering fuel and expending all solid rocket fuel before reentry.
Two liquid methane/oxygen (CH4/O2) fueled orbital maneuvering engines providing up to 2mt thrust each represent the only new technology required for the Aquila. A liquid methane/oxygen fuel engine provides fairly high specific impulse, a non-toxic propellant, and better storage (in terms of temperature, tank size and fuel boil-off) than a liquid hydrogen/oxygen solution. It provides the best compromise for extended operations and a basis for engines using Martian-derived fuel sources.
With help from the abort engines, the Aquila's total system maneuvering capacity is 1,150 m/s, more than enough to support braking and return from L1. For this, of course, the heat protective system needs to be certified for reentries at 40,000kph.
Thus, for the cost of developing a low thrust liquid methane-fueled engine and testing an improved high temperature reentry heat protection system, the space agency can assure crewed access to low earth orbit and extend it 5/6th of the way to the moon using existing launch vehicles. With adequate funding, developing and testing this vehicle within 6 years represents no great technological challenge.
Longer term access to space should focus on reducing costs per unit of mass launched. Once the IIS is complete, the shuttle will have no cargo delivery role, as better suited and cheaper expendable vehicles can fulfill this role for future projects. In the future, a reusable launch system is the only way to reduce costs, but earlier efforts, (including the X-30 and X-33 projects) have failed to produce viable vehicles. Given the contrasting success of the Air Force's EELV program that produced the Delta IV and Atlas V vehicles, a more gradual approach to full reusability seems more prudent than a crash program to develop a fully reusable vehicle. The fuel-mass ratio for single stage to orbit (SSTO) vehicles is too stringent for current material science and an air-breathing scramjet that could favorable alter this ratio will require years of development.
The most reasonable step to providing cheap access to space is to gradually develop a reusable system, and the best method for doing that is to start with the first stage of a launch vehicle, as the cheaper upper stage is often sent to geosynchronous transfer orbit (GTO) and beyond, making it unsuitable for recovery in any case. The following discussion again uses the Delta IV family (in this case the "Heavy") as a baseline, but does not preclude an Atlas-based vehicle, especially since the smaller mass and size of its kerosene-liquid oxygen tanks make it a viable reusable system.
The first stage of a two-stage launch vehicle needs to delver 4,000- 5,000 m/s gross delta v, a portion of which goes toward achieving altitude and overcoming gravity, not toward final orbital velocity. Even at sea-level engine performance, the mass fraction for this stage would leave between 25% and 33% of the vehicle for structure, second stage and payload. This would not prove a daunting technical challenge for a reusable vehicle.
A reusable first stage (RFS) (similar to proposed liquid return boosters for the shuttle) would return to a runway at the launch facility using jet engines. Using a lifting body design (like the X-33) and requiring only minimal reentry protection (the maximum velocity of the vehicle would probably not greatly exceed Mach 8) a five-engine "Heavy" reusable booster could launch the same payload as a three-engine "Heavy" expendable booster at perhaps five times the initial cost of the expendable, including research and development costs. Additional units would probably cost half as much. This price tag, in the billion dollar range spread over a few years, is fairly modest compared to estimates for the full Venture Star X-33 follow-on vehicle.
Assuming a market for ten heavy launches a year (and subsequent sections of this document would account for that number on their own), two producers of RFS launchers could be coaxed (with development and financial support similar to the EELV program) into developing a set of two or three of these launchers, lowering the overall costs to orbit, providing faster turnaround and assuring access to space. An RFS project could easily produce a viable vehicle by the end of the decade, replacing the launch capacity of the shuttle fleet and providing lower cost launch vehicles to support an L1/Lunar infrastructure (as detailed in Section III). Note that this project does not predispose any changes beyond a possible "stretch" fuel tank to existing upper stage designs, though a clean redesign or other alternative to the aging RL-10 family is probably already overdue.
The RFS vehicle can also form a test bed for more advanced technologies, perhaps replacing one engine with a scramjet to lower oxidizer mass and can eventually form the basis for a fully reusable vehicle, perhaps a mesh of the X-30 and X-33 design goals, by 2020.
NEXT: Section II
All pages and images ©1999 - 2005
by Geir Lanesskog, All Right Reserved