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29 July 2012
Added "Space Debris and Its Mitigation" to the archive.
16 July 2012
Space Future has been on something of a hiatus of late. With the concept of Space Tourism steadily increasing in acceptance, and the advances of commercial space, much of our purpose could be said to be achieved. But this industry is still nascent, and there's much to do. So...watch this space.
9 December 2010
Updated "What the Growth of a Space Tourism Industry Could Contribute to Employment, Economic Growth, Environmental Protection, Education, Culture and World Peace" to the 2009 revision.
7 December 2008
"What the Growth of a Space Tourism Industry Could Contribute to Employment, Economic Growth, Environmental Protection, Education, Culture and World Peace" is now the top entry on Space Future's Key Documents list.
30 November 2008
Added Lynx to the Vehicle Designs page.
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M Sarigul-Klijn & N Sarigul-Klijn, 2003, "Flight Mechanics of Manned Sub-Orbital Reusable Launch Vehicles with Recommendations for Launch and Recovery", AIAA 2003-0909. January 2003 (Revised April 2003).
Also downloadable from http://www.spacefuture.com/archive/flight mechanics of manned suborbital reusable launch vehicles with recommendations for launch and recovery.shtml

References and Referring Papers    Printable Version 
 Bibliographic Index
Flight Mechanics of Manned Sub-Orbital Reusable Launch Vehicles with Recommendations for Launch and Recovery
Marti Sarigul-Klijn Ph.D. and Nesrin Sarigul-Klijn*, Ph.D.
An overview of every significant method of launch and recovery for manned sub-orbital Reusable Launch Vehicles ( RLV) is presented here. We have categorized launch methods as vertical takeoff, horizontal takeoff, and air launch. Recovery methods are categorized as wings, aerodynamic decelerators, rockets, and rotors. We conclude that both vertical takeoff and some air launch methods are viable means of attaining sub-orbital altitudes and wings and aerodynamic decelerators are viable methods for recovery. These conclusions are based on statistical methods using historical data coupled with time-stepped integration of the trajectory equations of motion. Based on the additional factors of safety, customer acceptance, and affordability, we also conclude that the preferred architecture for a manned sub-orbital RLV is Vertical Takeoff using hybrid rocket motor propulsion and winged un-powered Horizontal Landing onto a runway ( VTHL).
NOMENCLATURE
BC = Ballistic Coefficient
g = Earth's gravitational acceleration
go = Gravitational constant
h = Peak trajectory altitude
Isp = Specific Impulse
M = Mach number
MR = Propellant mass ratio, Wo/Wf
to = Engine burn time
T/W = Thrust to weight ratio
T = Thrust
Wf = Burnout weight
Wo = Takeoff weight
Wp = Propellant weight
ΔV = Change in velocity (delta V)
δp = Propellant mass fraction
INTRODUCTION

Space tourism is predicted to be a multibilliondollar industry when a safe and economical space vehicle is built.[1] With the exception of NASA's Space Shuttle and the Russian Soyuz, there are no other manned vehicles that can leave the Earth's atmosphere and return safely. The shuttle is expensive and current US law prevents tourists from buying a ride, even if they could afford to. A tourist seat is currently available on the Russian Soyuz every 6 months, but at a cost of $20 million.

The St. Louis based X-Prize Foundation is offering a $10 million prize to the first team that launches a privately financed vehicle capable of carrying three people to a 100 kilometer (328,000 feet) sub-orbital altitude and repeat the flight within two weeks. Only one person and equivalent ballast substituted for the other two persons are actually required to make the flights. Also no more than 10% of the vehicle's first flight non-propellant mass may be replaced between flights. The XPrize is being offered to help speed the development of concepts that will reduce the cost of manned space flight.[2]

However, there exists considerable disagreement on the correct architecture for a safe and economical space vehicle. Fundamentally, teams must decide on a takeoff mode and on a landing mode.

There are three basic takeoff modes available; vertical, horizontal, and air launch. Early success in manned space flight was achieved with vertical takeoff using rocket power with such vehicles as the Russian Vostok and American MercuryRedstone rockets. Sub-orbital manned space flight was achieved using air launch takeoff with the American rocket powered X-15. So far no suborbital or orbital flights have been achieved with either rocket powered horizontal takeoff aircraft such as the German Me-163 (peak altitude of 42,000 ft or 12,800 m) or combined rocket and jet powered horizontal takeoff aircraft such as the Lockheed NF-104 (peak altitude of 120,800 ft or 36,800 m) and the British Saunders-Roe SR-53 (peak altitude of 67,000 ft or 20,400 m).

In terms of landing mode, there are 4 major recovery methods available: wings, aerodynamic decelerators (such as parachutes), rockets, and rotors. Wings and parafoils (a rectangular ram-air lifting parachute) are considered horizontal landing. The rest are considered vertical landing.

Horizontal landing with wings and wheeled landing gear has been successful for manned piloted vehicles. Examples include the American X-1, X-2, X-15, and Space Shuttle.

Vertical landing using parachutes into the water has also been successfully demonstrated in the Mercury, Gemini, and Apollo programs, although one astronaut did almost drown during the second manned Mercury landing. The Russians have had success with parachute landings using a retrorocket to soften the landing. However, there was one cosmonaut fatality during the first manned Soyuz due to a failure of its parachute system.

Other landing methods have not been used for manned earth return missions, although some tests have been conducted. These include vertical landing using rockets only (such as the McDonnell DC-X), vertical landing using a rotor ( Rotary Rocket Company's Roton), vertical landing with parachutes using airbags to soften the landing (Kistler K-1), and horizontal landing using a parafoil parachute (NASA's X-38). Manned vertical landing using rockets has been successfully demonstrated on the Moon during the Apollo program.

The 3 takeoff modes and 4 landing modes result in 12 possible architectures for a sub-orbital RLV. As this paper will show, only 4 of these 12 architectures are viable candidates for the X-Prize.

ASCENT FLIGHT MECHANICS

The adjacent figure presents the typical trajectory of a X-Prize vehicle, a trajectory that is also considered typical for vehicles planned for the sub-orbital tourist market. A substantial portion of the trajectory is above the atmosphere. The only propulsion system that can work in the vacuum above the atmosphere is a rocket engine.

Specific Impulse and the Rocket Equation

A quantity called specific impulse and abbreviated by "Isp" provides a convenient measure of a rocket engine's intrinsic efficiency. The specific impulse or Isp of a rocket and propellant combination is analogous to "miles per gallon" for an automobile. The Isp of a rocket-propellant combination is defined as the number of seconds a pound (lb) of propellant will produce a lb of thrust.

Typical X-Prize® Trajectory[2]

For example, an Isp of 200 seconds means that a rocket engine would consume 1 lb (or kilogram) of propellant when producing 1 lb (or kilogram force) of thrust for 200 seconds. Once a fuel and oxidizer are chosen, the Isp is largely determined by the energy contained in its propellants. Generally speaking, designers strive for the highest Isp they can achieve.

Realistic Isp for a space tourist vehicle's engine is in the order of 200 to 275 seconds based on today's technology. Many X-Prize concept designs are erroneously based on much higher Isp. Rocket engine books and reference sources provide convenient tables of Isp for various propellant combinations. These tables often give data for the theoretical operation of rocket engines using a combustion chamber pressure of 1000 pounds per square inch (psi or 68 bar) operating in a vacuum. Rocket engines have lower Isp when operating in the atmosphere or at lower combustion chamber pressures. For example, the theoretical vacuum performance of a kerosene and liquid oxygen ( LOX) engine is about 350 seconds. In contrast the demonstrated Isp at sea level for the Saturn V's F-1 engine is 265 seconds, the Delta's RS-27 is 262 seconds, and the Atlas's MA-5 is 259 seconds. Tourist vehicles will likely have engines that operate at much lower combustion chamber pressures than these engines since most will be pressure fed instead of turbo pump fed. For example, sea level Isp for a pressure fed keroseneLOX engine with a 250 psi (17 bar) chamber pressure is about 225 seconds.

Once Isp is known, then the ideal rocket equation, which is also called Tsiolkovsky's equation in honor of the Russian schoolteacher who first derived it over 70 years ago, indicates the maximum velocity that can be obtained from a load of propellant. It is given by the following equation:

ΔV = Isp go ln(Wo/Wf)     (1)

It says that the change in velocity (called delta V and symbolized by ΔV) is directly proportional to the specific impulse, Isp, multiplied by the gravitational constant, go (equals 32.174 ft/sec2 or 9.806 m/sec2), and multiplied by the natural logarithm of the ratio of the weight of the rocket at takeoff, Wo, and the weight of the rocket at burnout, Wf. This ratio of takeoff weight to burnout weight, Wo/Wf, is called the propellant mass ratio and is sometimes written as MR. It has a value always greater than 1.

When comparing design concepts, it is much easier to use propellant mass fraction, δp, which is related to propellant mass ratio, MR, by:

MR = 1/(1-δp)     (2)

Propellant mass fraction, δp, is defined as the amount of propellant relative to the takeoff weight. For example, if the mass of a rocket at takeoff is 50% propellant, then its propellant mass fraction, δp, is 0.50. In this example the propellant mass ratio, MR, would be 2 since the takeoff weight is twice the burnout weight.

Hence if we know the propellant mass fraction, δp, and its specific impulse, Isp, we can then determine a rocket's delta V, ΔV.

Delta V Budget for Sub-Orbital Flight

The next question that we need to answer is "How much delta V is needed to reach a suborbital height of 100 kilometers?" The smallest delta V to reach 100 km (328,000 ft) would be achieved by firing the launch vehicle out of a large cannon vertically. In the absence of an atmosphere, the cannon would have to accelerate the launch vehicle to about 4,600 feet per second (fps or 1,400 m/s) to reach 100 km above the earth's surface. Placing this hypothetical cannon at 25,000 ft (7,600 m) altitude to represent an air launch would reduce the ideal delta V somewhat to about 4,400 fps (1,340 m/s). However, the delta V that a launch vehicle's rocket motor must actually provide is greater than this amount because of several losses.

The first loss is known as the gravity loss. Gravity loss arises because part of the rocket engine's energy is wasted holding the vehicle against the pull of Earth's gravity (g). Gravity losses depend on the takeoff thrust to weight (T/W) ratio. A T/W ratio of 1.0 means that the rocket engine's thrust just equals the vehicle's weight. The T/W ratio obviously needs to be greater than unity in order to climb. For launch vehicles designed to launch payloads into low earth orbit ( LEO), initial T/W is in the order of 1.2 to 1.5 for liquid rockets and somewhat higher for solid rockets. Gravity losses are in the order of 2,000 to 4,000 fps (600 to 1,200 m/s) for a vertical trajectory to 100 km (328,000 ft).

Drag is another loss and is caused by friction between the launch vehicle and the atmosphere. Drag losses are in the order of about 500 fps (150 m/s) for medium sized launch vehicles such as the Delta or Atlas rockets for an earth to orbit trajectory. A long slender cylinder with a pointed nose is a favored shape to reduce drag losses since over three-quarter of drag losses are caused by supersonic drag. Also drag losses are subjected to the "cubed-squared" law. As an object's external dimensions increase, the surface area increases with the square of the dimension while the volume increases with the cube. Since drag is a function of surface area and not volume, then increasing the launch vehicle size will reduce drag losses. For example, the huge Saturn V moon rocket had drag losses of only 130 fps (40 m/s). The cubed-squared law is the reason why launch vehicles launched from the earth's surface should be at least 30 to 40 ft (9 to 12 m) long in order to leave the earth's atmosphere. Otherwise drag losses are simply too high.

Steering loss is the last major loss and is caused by the need to steer the launch vehicle. It arises because the instantaneous thrust vector is not always parallel to the current velocity vector. This mismatch is necessary; otherwise, it would be impossible to steer a launch vehicle. In order to get an appreciation for steering losses, the Delta II rocket has only 110 fps (34 m/s) of steering losses as compared to the Space Shuttle's 1,200 fps (365 m/s), both for a low earth orbit trajectory.[3] Vertical takeoff vehicles have much less steering losses as compared to horizontal takeoff vehicles since a horizontal vehicle must pitch upward 90 degrees to change its direction of flight to vertical.

Notice that a compromise flight profile must be chosen to minimize losses. A high thrust to weight (T/W) ratio would reduce gravity losses but it would increase drag losses since the vehicle would be traveling very fast while low in the atmosphere. On the other hand, a low T/W ratio would reduce drag losses but greatly increase gravity losses.

Flight trajectory simulation computer programs, such as NASA's POST (Program to Optimize Simulated Trajectories) are needed to numerically integrate the equations of motion and to find the best trajectory. These equations typically do not have a closed-form solution because of the complex nature of the drag and steering losses. The amount of input and output data for various flight simulation programs can range from the simple to vast; for example, the POST program requires hundreds of inputs.

An approximate delta V to reach 100 km is 7,000 to 8,000 fps (2,100 to 2,400 m/s) for vertical takeoff, with slightly less delta V needed for air launch, and significantly more required for horizontal takeoff.

A closed form equation can only be obtained if the steering and drag losses are not included. The following equation gives the peak trajectory altitude, h, which is the height a rocket reaches in a vertical trajectory in the presence of gravity:

h = ½ Isp2 go ln MR (ln MR - 1/T/W + 1/(MR × T/W))     (3)

This equation is plotted in the next figure for two different values of Isp and for a T/W ratio of 1.5.

These plots assume no aerodynamic drag or steering losses. Height corrected for steering and drag losses would be 70% to 80% of that shown. The figure shows that a sub-orbital vehicle needs a propellant mass fraction, δp, of 0.512 with an Isp of 275 seconds or a δp of 0.574 with an Isp of 225 seconds in order to reach a sub-orbital altitude of 100 km (328,000 ft). This assumes a launch from the ground and ignores aerodynamic drag and steering losses.

The next figure presents the results from a trajectory program that accounts for drag and steering losses. The blue solid line is the altitude in feet divided by 100,000, the red dashed line is acceleration in g's, and the green dotted line is Mach number. Peak aerodynamic pressure for this particular trajectory during both ascent and reentry is 750 pounds per square foot (psf), equal to 470 knots equivalent airspeed (KEAS). Further results from this trajectory program show that the propellant mass fraction needs to be about 0.63 for a vertical launch from the ground for a vehicle that is over 50 ft (15 m) long, assuming an average Isp of 225 seconds.

In addition to the propellant that is burnt by the rocket engines, all launch vehicles carry residual and reserve propellants. Residual propellant is propellant that is trapped in the propellant tanks or lines. Reserve propellant is extra propellant loaded onboard the vehicle to ensure that enough is available to complete the mission. Combined residual and reserve propellant mass fraction can be as little as 0.01 for a turbo pump fed liquid fueled engine with a mixture control system to as much as 0.07 for a simple pressure fed liquid rocket or for some solid or hybrid rocket motors. Residual and reserve propellant must be added to the propellant determined by a trajectory program to get the propellant mass that must be actually carried by the launch vehicle. Hence total propellant mass fraction, δp, can be as much as 0.70 for a vertical takeoff vehicle in order to reach 100 km (328,000 ft).

Estimation of Propellant Mass Fractions

Aerospace vehicle propellant mass fraction estimation is evolutionary rather than revolutionary; that is, a new vehicle design's propellant mass fraction is usually an evolutionary change from previously existing designs. Hence they can be estimated statistically from historical trends.

The next figure presents the propellant mass fractions, δp, for various types of vehicles. Vertical takeoff rockets are represented with red triangles. These are multistage expendable vehicles and only a portion is safely returned to earth. Air launched rocket powered vehicles are shown as blue diamonds and, except for Orbital Science's Pegasus, the entire vehicle returns to earth in one piece. Supersonic jets are shown for reference as green squares. Finally, combined rocket and jet powered planes are shown as purple circles.

Notice that the type of vehicle has a strong effect, with horizontal takeoff using combined jet and rocket power having the lowest rocket propellant mass fraction (0.09 to 0.13), and vertical takeoff rockets having the highest (0.85 to 0.94). Also notice that propellant mass fractions increase with increasing takeoff weight. The trends for each type of vehicle fall along approximately straight lines. Finally, reusable vehicles have a lower propellant mass fraction as compared to vehicles that have expendable parts.

Historically, all vehicle designers were highly motivated to increase propellant mass fraction as much as possible, since doing so provides both tremendous commercial and military advantages. Hence, the propellant mass fractions for new space tourist vehicles should follow historical trends.

TAKEOFF MODE TRADE
Vertical Takeoff

The Mercury-Redstone that carried American astronauts on their first sub-orbital flights is an example of a vertical takeoff vehicle. A vertical takeoff rocket can easily carry the necessary propellant mass fraction required to reach 100 km (328,000 ft). Vertical takeoff vehicles also have significant margin that allow them to use relatively low technology pressure fed rocket motors. Vertical takeoff rockets are basically pressurized balloons under an axial compression load and are subject to very little bending and no twisting moments during takeoff. Unlike horizontal takeoff vehicles, precise longitudinal control of center of gravity is not required which eliminates the need for multiple bulkheads inside their fuel and oxidizer tanks and eliminates the need for intertank propellant transfer pumps.

Mercury-Redstone "Milk-Stool" takeoff pad

Vertical takeoff requires a prepared takeoff site. A portable "milk stool" type takeoff pad should be sufficient for a sub-orbital tourist vehicle. Launch Propellant Mass Fraction preparation time does not need to be excessive. During the Persian Gulf War the Iraqi were able to erect, fuel and launch their Scud missiles within 1 hour of arriving to a bare concrete pad. A 3-person, 100 km and Mach 3.5 capable space tourist vehicle should be smaller than a MercuryRedstone, which was capable of 187 km (613,000 ft) altitude and Mach 7. Hence a milk-stool takeoff pad should be smaller than the one shown in the picture.

Vertical takeoff has been picked by every space transportation study during the past 20 years as the preferred takeoff method for an earth to orbit RLV. These studies include the European Space Agency's (ESA) Future European Space Transportation Investigations Program (FESTIP), Russian's Oryol program, Japan's HOPE, and the United States' Access to Space Studies, Space Transportation Architecture Studies, and most currently the US Space Launch Initiative (SLI).[4]

Horizontal Takeoff

Approximately half of the X-Prize concepts listed on the X-Prize web site are horizontal takeoff. Based on this study, horizontal takeoff concepts will be unable to reach the X-Prize altitude goal of 100 km (328,000 ft) using current technology propulsion and structural materials.

Propellant mass fraction is much less for horizontal takeoff than for vertical takeoff vehicles. In addition to propellant tanks and rocket engines, many concepts have large wings and heavy landing gear sized for their fully fueled takeoff weight. In general, horizontal takeoff aircraft are subject to high bending and twisting moments as compared to vertical takeoff vehicles. Lightweight aircraft are quite flexible, especially in fuselage longitudinal bending, wing spanwise bending, and wing torsional bending. These have a major effect upon stability characteristics that in turn affects structural weights. For example, a typical sweep wing transport at high subsonic speeds will have a reduction in elevator effectiveness of about 50% due to fuselage flexibility effects. Aileron effectiveness is reduced by 50% to 100% because deflecting the ailerons twist the wing in the opposite direction causing a condition known as "aileron reversal." Thus many aircraft components are not designed for strength, but for stiffness, which greatly increases structural weight.

Furthermore, precise control of longitudinal center of gravity (CG) is required to keep the CG within a narrow longitudinal range, otherwise flight control will be lost. This may necessitate multiple fuel tanks and fuel transfer pumps to manage the fuel burn. This adds more weight.

There are two types of horizontal takeoff vehicles:

  1. Rocket only power
  2. Combined jet and rocket power.
Rocket Powered Horizontal Takeoff.
There have been several horizontal takeoff rocket powered aircraft built. These include the German Me-163, the Russian Mikoyan I-270, and recently the XCOR's EZ-Rocket. During takeoff, these aircraft spend a considerable time firing their rocket engines in the horizontal direction consuming propellant; this is another reason why they are not capable of high peak altitudes. The Me-163 was the highest flying of these aircraft and had a demonstrated altitude of only 42,000 ft (12,800 m), although it had a calculated peak altitude of 52,000 ft (15,800 m).[5] The Bell X-1 was normally air launched, but Chuck Yeager flew it once from a runway to an altitude of 23,000 ft (7,000 m).

Rocket engine thrust can be less than the initial takeoff weight, i.e., T/W can be less than 1.0, in a horizontal takeoff vehicle. Typically, rocket engines weigh about 1% to 2% of the takeoff weight. In contrast wings and wheels weigh about 20% of takeoff weight for a horizontal launch vehicle. So although a rocket powered horizontal takeoff vehicle benefits from a small reduction in the size of its rocket engine, it does so at the penalty of a large increase in its overall empty weight due to its larger wings and landing gear.

Messerschmitt Me-163B (USAF Museum image)

A common misconception is that a rocket powered vehicle using horizontal takeoff will use less propellant to climb to an altitude as compared to a vertical takeoff vehicle. Propellant mass, Wp, is give by rocket engine thrust, T, multiplied by engine run time, tb, and divided by specific impulse, Isp:

Wp = T tb/Isp     (4)

Although a horizontal takeoff vehicle's thrust is less than a vertical takeoff's, the engine run time is longer since the horizontal takeoff vehicle climbs along an inclined flight path. Hence if flight velocities are the same and drag is ignored, then propellant burnt is the same. In reality, a horizontal takeoff vehicle will use more propellant since it must overcome aerodynamic drag for a longer time and it must overcome the additional drag caused by its wing producing lift (known as induced drag). Another way to look at this is that the Work (force multiplied by distance) done against gravity is the same for an object raised vertically as it is for an object raised along an inclined path if friction is ignored. If friction is included, then the shortest path (vertical) requires less Work.

Even if a rocket powered horizontal takeoff vehicle conducts a sharp pull-up to a vertical trajectory immediately after takeoff, it will use 20% to 30% more propellant as compared to a vertical takeoff vehicle because the horizontal takeoff vehicle must fire its rocket motor for 20 to 30 additional seconds.

Horizontal Takeoff Combined Power.

Surprisingly many of the proposed X-Prize vehicles are horizontal takeoff concepts using combined jet and rocket power. Jet engines are very large in comparison to rocket engines and they weigh 10 to 20 times more than a rocket engine of similar thrust. On the other hand, a jet engine's specific impulse, if considered on a thrust per unit flow of propellant carried, i.e. kerosene only, is some 20 times higher than that of a rocket. This is why jet engines are preferable for cruise missiles and aircraft. Finally, a jet engines thrust decreases with altitude at subsonic flight speeds; approximately half of sea level thrust at 22,000 ft (6,700 m) and quarter at 40,000 ft (12,000 m), while a rocket engine's thrust increases by 15% to 25% in a vacuum as compared to sea level.

A combined powered vehicle must carry both jet engines sized to provide all the thrust needed to fly the aircraft and rocket engines also sized for the same task. The two propulsion systems not only add weight, but also use up internal volume that would otherwise be used to carry either propellant or payload. Propellant tanks must be divided to carry both jet fuel and rocket propellants. It is not surprising that these designs have the smallest propellant mass fraction of all the various concepts.

Although not a combined powered vehicle, the Concorde's propellant mass fraction of 0.51 probably represents the maximum internal fuel load possible for a supersonic horizontal takeoff concept. The designers of the Concorde had to build three different full-scale prototypes before they were able to optimize the Concorde design so that it could carry enough fuel to cross the Atlantic. Furthermore, in order to carry enough fuel, the Concorde passengers and payload had to be reduced to only 7% of its takeoff weight as compare to 23% for the subsonic Boeing 747. It is very unlikely that any combined powered RLV designer can improve upon the Concorde's propellant mass fraction, especially considering that the Concorde carried the majority of its fuel in its wings. In contrast, a RLV will have to carry the majority of its propellant in the fuselage because a rocket engine's turbo pumps need an inlet pressure of about 40 psi (2.7 bar) to prevent pump cavitations; a pressure far too high for wing tanks. In case of a simpler pressure fed rocket, the tank pressures are greater than 250 psi (17 bar).

At least four different combined jet and rocket powered aircraft were built and flown. Of the four, the Lockheed NF-104 was the only one that could fly supersonic on the thrust of its jet engine alone. In 1959 Air Force Captain Joe Jordan set an altitude record of 103,000 ft (31,400 m) in a standard F-104 using jet power only. In 1963 the NF-104's rocket engine allowed Major Robert Smith to zoom climb to a record 120,800 ft (36,830 m), only 17,800 ft (5,430 m) higher. Of the three NF-104's built, one was lost to a rocket engine explosion and another was lost to an out of control flight accident. The last one is on a pole in front of the US Air Force's Test Pilot School at Edwards AFB. The NF-104's flew 302 flights accumulating 8.6 hours of rocket engine operation.[6]

Lockheed NF-104 (USAF image)

The remaining three combined powered aircraft were capable of only subsonic flight when powered by their jet engines. In this respect, they are similar to all of the currently proposed X-Prize concepts in that none of these X-Prize concepts are expected to be capable of supersonic flight using their jet engines.

The Republic XF-91 Thunderceptor was first flown on jet power on 9 May 1949. The first rocket-powered flight was in fact unplanned and came on 11 September 1952 when the Thunderceptor's J47 turbojet flamed out on takeoff. Of the two built, one was destroyed when its rocket engine exploded, nearly blowing the tail off the aircraft. The other aircraft is on display at the Wright-Patterson AFB Museum. Peak altitude obtained with both rocket and jet power was 48,000 ft (14,600 m) and maximum flying time on jet only power was 25 minutes.[7]

Republic XF-91 Thunderceptor (USAF image)

The British Saunders-Roe company built two SR-53 combined powered aircraft. One was damaged to a rocket engine explosion during a ground test in 1955. The SR-53 first flew on 16 May 1957. The test program was stopped when a SR-53 crashed and fatally injured its pilot during takeoff on 5 June 1958. Peak altitude with both rocket and jet power was 67,000 ft (20,400 m).[8]

Saunders-Roe SR-53

Finally, France's Sud Aviation built eight Trident prototypes of 3 different designs. First flight was on 19 July 1955 and they logged more than 600 flights including 220 on rocket power. One Trident was lost on its second flight and another crashed during the 1957 Paris air show. They were capable of zooming to 79,500 ft (24,200 m) when both jet and rocket engines were operating.[9]

Although almost half of the X-Prize entrants have chosen combined powered horizontal takeoff concepts, these historical examples demonstrate that they will have a difficult time reaching 100,000 ft (30,400 m), much less the X-Prize goal of 328,000 ft (100,000 m).

Sud Aviation SO 9050 Trident
Air Launch

In an attempt to solve some of the problems of horizontal takeoff, several air launched methods have been proposed. Air launched methods can be categorized into six methods:[10]

  1. Captive on top
  2. Captive on bottom
  3. Towed
  4. Aerial refueled
  5. Internally carried
  6. Balloon

Air launching reduces the propellant mass fraction required to reach 100 km as compared to a vertical ground launch. For example, the X-15 carried a propellant mass fraction, δp, of 0.55 and was capable of sub-orbital flight. "Zoom" launching a RLV with the carrier aircraft to a flight path angle of 50 to 60 degrees above to the horizon reduces the δp needed to reach 100 km to about 0.5. A RLV with a δp of 0.5 must carry its empty weight in propellant. In contrast, a ground launched vertical takeoff RLV with a δp of 0.63 must carry twice its empty weight in propellant. Hence air launching can reduce RLV size.

The reduction in δp is due to shorter climb distance (85 km to 94 km instead of 100 km) and increased rocket engine specific impulse. Specific impulse, Isp, is improved at altitude because of better thrust expansion in the engine nozzle and to using a larger nozzle properly sized for the launch altitude.

Air launching also reduces the aft fuselage damage caused by acoustic energy from the engine since there is no reflection from the ground during launch and local air density is lower at launch.

Furthermore, air launch RLVs can operate with minimum launch site requirements. A takeoff pad is not required.

On the other hand, winged RLVs that are air launched may require precise control of longitudinal center of gravity. This means that fuel and oxidizer tanks may have to be divided into smaller tanks with additional tank bulkheads and transfer pumps. Both the X-15 and X-34 had propellant tanks subdivided into multiple tanks, but clever design can limit the number of tanks.

Air launching can either reduce or increase the peak aerodynamic pressure that the RLV is subjected to. Peak dynamic pressure is one of the elements that determine vehicle structural design. For example, the Orbital Science's air launched Pegasus XL experiences over 1,250 pounds per square foot (psf or 0.6 bar) aerodynamic pressure, twice the Space Shuttle's, even though the Pegasus is launched at 38,000 ft (11,500 m). On the other hand, zoom launching a RLV with the carrier aircraft, launching from a slow flying carrier aircraft, or balloon launching can reduce RLV peak dynamic pressure to as little as 1/3 of a vertical ground launch or to about 200 to 300 psf.

Air launched winged RLVs are subjected fuselage bending loads similar to horizontal takeoff vehicles. This means that their empty weight fractions are greater than vertical takeoff vehicles.

Air launching requires a specially modified carrier aircraft capable of carrying the RLV. The carrier aircraft will typically be heavier than the RLV itself; for example, the 480,000 lb (220,000 kg) B-52 carrier aircraft carrying a 33,000 lb (15,000 kg) X-15. This limits the size of the RLV and means that there is limited growth potential for air launching.

In some air launch concepts, parts (such as explosive bolt debris or cables) are designed to or have the potential to fall off. For this reason, government and safety regulators may require that an air launch be conducted over certain areas such as a government range or over the open ocean. This potentially can negate one of air launching's primary advantages, which is the ability to operate free of national range scheduling constraints.

Propellant boil off can be a major problem for those concepts that both use cryogenic propellants and have the RLV carried outside the carrier aircraft. Propellants are heated by radiation heating from the sun and convective heating from the air stream. For example, the X-15 boiled off 60% to 80% of its liquid oxygen ( LOX) during its 45 minute to 1 hour climb and ferry while attached to its B-52 carrier aircraft. The LOX was replenished from an internal insulated tank carried in the B-52's bomb bay.

Finally air launched RLV engines are typically started after they are dropped from the carrier aircraft for the safety of the carrier aircraft. If the rocket engine fails to start, then propellant must be dumped quickly and an emergency landing must be completed. This happened several times during the X-15 rocket plane's 199 flights.[11]

Captive on top

No examples of captive on top RLVs have been actually built, but the Space Shuttle's approach and landing demonstrator, the Enterprise, used this method to test its landing. Advantages of this method include the capability to carry a large RLV on top of the carrier aircraft and lighter RLV landing gear sized only for landing. Disadvantages include penetrations on the windward side of the RLV's thermal protection system (TPS) for attachment hard points and extensive modifications (high cost) to the carrier aircraft. Also the RLV must have active controls at release from the carrier aircraft to prevent it from hitting the carrier aircraft. Since the RLV's wings must be large enough to support the RLV at separation from the carrier aircraft, the RLV itself may have to be a multi-stage vehicle and its wings may have to be released and staged after an aerodynamic pull-up into a vertical trajectory.

Placing a RLV on top of the carrier aircraft destroys the lift produced by the fuselage and causes a large amount of drag that in turn limits launch altitude. For example during its approach and landing test flights, the Space Shuttle was launched at altitudes between 19,000 to 26,000 ft (5,800 to 8,000 m) from its carrier Boeing 747, even though a clean 747 normally cruises at 38,000 to 45,000 ft (11,500 to 13,700 m).

X-15 air launching from B-52 (NASA image)
Captive on bottom.

Examples include the X-15 and X-34. Advantages include proven and easy separation from the carrier aircraft, leeward side penetrations and hard points on the RLV that eliminate some TPS concerns, lighter RLV landing gear sized for landing, and the option of sizing the wing smaller than required for level flight at the release altitude and airspeed. Disadvantages include limits to RLV size due to under the carrier aircraft clearance limitations and the high cost of carrier modifications. A new design carrier aircraft with tall landing gear can eliminate the clearance limitations.

Towed.
One of the first occurrences of towing a rocket-powered aircraft was during the summer months of 1942 at Peenemünde, Germany. Twin engined Bf-110C fighters were used to tow prototypes of the Me-163 rocket fighter for flight tests, typically to altitudes of 16,400 ft (5,000 m).

The primary advantages of a towed air launch are easy separation from the towing aircraft and low cost modifications to the towing aircraft. Safety concerns include broken towlines and a towing aircraft takeoff abort.

The principal disadvantage of towing is that the RLV's wings and wheels must be sized for takeoff with a full propellant load. Towing provides some improvement as compared to rocket powered horizontal takeoff, but in order to reach 100 km, a multi-stage RLV may be needed in which the wings and wheels are staged after the tow-line release and completion of the aerodynamic pull-up maneuver.

Aerial Refueled.

Aerial refueling is only proven for kerosene fuels. Cryogenic propellants like liquid oxygen have the potential of freezing the refueling probe to the refueling line. The principal advantage of aerial refueling is that it reduces the size of the RLV landing gear and wing. Note that aerial refueling does not reduce the size of the jet engines; they must be sized to maintain level flight for the fully fueled RLV, and aerial refueling does not reduce the strength of the wings, which must be strong enough to support the fully fueled RLV weight. Aerial refueling provides some improvement to a combined powered horizontal takeoff concept, but in order to reach 100 km, a multi-stage RLV may be needed in which the wings, jet engines, and landing gear are staged after the completion of a pull up maneuver.

Internally Carried.

Internal air launch has been demonstrated before. On 24 October 1974 a C-5A Galaxy dropped a 78,000 lb (35,500 kg) LGM-30A Minuteman I missile using drogue chutes to extract the missile and its 8,000 lb (3,600 kg) launch sled. The missile was then successfully fired.

Advantages of internally carried concepts include little or no modification to the carrier aircraft. Most propellant boil-off concerns are eliminated since the RLV is not subject to either radiation heating or convective heating. The RLV is in a benign environment inside the carrier aircraft that allows maintenance and safety problems to be detected prior to launch. Release altitude can be at a higher altitude because the RLV does not increase the carrier aircraft's drag.

Minuteman launching form C-5A (USAF image)

The main disadvantage is that the RLV must be sized to fit inside the carrier aircraft. Also operations must be conducted over the water since many parts such as a launch sled fall away.

Balloon Launch.

Balloon launch requires operating a very large balloon. Launch can occur only on calm days. Since the balloon comes back unmanned, the potential to damage either the balloon or things on the ground is high.

ATMOSPHERIC ENTRY

The next figure shows the Mach number and deceleration experienced by an object falling vertically. The figure was obtained by numerical time stepping of the equations of motion for a blunt object falling from 328,000 ft (100 km). The solid blue line represents Mach number, which peaks at 3.5 at 100,000 ft (30,500 m). The dashed red line represents deceleration, which peaks at 5.5 times the earth's gravity at 70,000 ft (21,300 m).

For objects entering the earth's atmosphere, peak entry deceleration is a function of the trigonometric sine of the entry angle and the square of the entry airspeed. Since a X-Prize suborbital trajectory is mostly vertical in both the up and down directions, then entry deceleration is a function of entry airspeed only, which in turn is determined by the trajectory's peak altitude (h).

The next figure shows the peak entry deceleration for two different ballistic coefficients (BC). BC is the ratio of the mass of a vehicle divided by the product of drag coefficient and drag area. Since drag coefficient varies with Mach number, then BC can be expected to vary with Mach number as well. A BC of 80 pounds per square foot (psf or 0.038 bar) is representative of the Mercury capsule whereas a BC of 160 psf (0.076 bar) is the Soyuz descent capsule. Notice that the X-Prize Foundation's altitude goal of 328,000 ft (100 km) is about the highest altitude that a tourist can be expected to tolerate comfortably. Also notice that peak deceleration does not vary too much with BC.

Reentry aerodynamic pressure is set by BC and the peak height of 100 km. The next figure shows the peak aerodynamic pressure in terms of equivalent airspeed in knots (KEAS). Peak dynamic pressure occurs between Mach 2 and 2.8 depending in BC. BC must be kept very low in order to use general aviation like structures. Otherwise stiffer structures similar to that used in fighter aircraft are needed.

Total temperature can momentary be expected to be approximately 1000 degree Fahrenheit (540 C) due to the Mach 3+ entry. The actual equilibrium surface temperature will depend on the heat sink characteristics of the vehicle structure. High temperatures will eliminate the large general aviation like Plexiglas windows seen in many Xprize concepts. Instead double pane high temperature windows may be required as well as some minimal thermal protection for the rest of the vehicle.

LANDING MODE TRADE

After a RLV decelerates to subsonic speeds, a landing mode must be selected. There are four major landing modes available:

  1. Wings
  2. Aerodynamic decelerators
  3. Rockets
  4. Rotors

In addition, a designer has the choice of landing the vehicle on its tail (tail sitter) or landing it on its side (horizontal lander). Tail sitters have the advantage of having only one structural load path; the vehicle can be designed mostly to axial loads. This saves weight and is the main reason that tail sitters have been considered in reusable Single Stage To Orbit ( SSTO) vehicles studies, since weight is critical for such vehicles. A tail sitter design is not required for either a sub-orbital or two stage to orbit vehicle.

A disadvantage of a tail sitter is that they may fall over upon landing, unless the landing area is perfectly flat and the winds are low.

Wings

Horizontal landing using wings and wheeled landing gear has been consistently successful, well proven, and has been used by the American X-1, X-2, X-15, and Space Shuttle, and Russian Buran space vehicles. It provides the softest landing of all the landing methods with landing touchdown vertical velocities averaging about 2 feet per second (fps or less than 1 m/s).

Wings and wheels can be sized to land even the largest RLV. For a vertical takeoff vehicle the wings can be sized for landing weight only, which significantly reduces their size and drag. For example, the left side of the next figure shows the wing size needed to land at 195 knots (100 m/s) an empty X-34 that weighs 18,000 lbs (8,200 kg). The right side of the figure shows the wing needed to horizontally take-off at the same airspeed a fully fueled X-34 weighing 48,000 lbs (21,800 kg).

In addition, a vertical takeoff RLV's wing can use blunt leading edges that reduce thermal protection problems. Sharp leading edges provide much better supersonic lift versus drag performance and are needed in horizontal takeoff RLVs. However sharp leading edges will get much hotter than a blunt leading edge. Heating rates are a square root function of leading edge radius.

Wings and wheels require long runways and either a highly skilled and proficient human pilot or a complex automated landing system. Based on historical data, wings and wheels will constitute about 20% of the landing weight. Although the ascent drag of the wings detracts from ascent performance, this turns out to be a minor loss.[3]

Critical to the design of a winged vehicle is the need for large dive brakes or some other means to minimize pullout acceleration g's because of the vertical nature of the entry trajectory. Pullout acceleration g's can be minimized by first allowing the vehicle to decelerate to subsonic speeds before transitioning from a vertical descent to a horizontal glide.

Aerodynamic Decelerators

There are three types of aerodynamics decelerators.[12] These are:

  1. Parachutes
  2. Parafoils
  3. Rigid or semi-rigid decelerator

Aerodynamic decelerators do not need a long runway and a highly trained human pilot or a complex automated landing system. However, the effort to develop a reliable man-rated system is illustrated by the X-38's four-year flight test program which consisted of 26 drops using 3 different size parafoils and 4 different test vehicles dropped from C-130 and B-52 aircraft.[13]

Parachutes

Parachutes are limited to landing weights less than about 40,000 lbs (12,200 kg). A parachute landing system can be the lightest weight recovery system with the parachute portion equal to only 2.8% to 6% of the landing weight.

Parachute landing needs a large flat landing area since there is little control over the precise landing spot unless a controllable parachute is used. The US Army has been successfully experimenting with a controllable drogue chute that allows relatively precise landings with a 330 ft (100 m) circular error probable (CEP), but the system is not operational yet[15]

The main challenge with parachute system is in the attenuation of the landing impact. In many concepts, the crew compartment separates from the rest of the vehicle to allow the use of a better landing attenuator for the crew. There are five attenuation methods:

Water landing. Parachute landing into the water has been proven to be the lightest weight recovery system of all landing systems. This method was successfully used in the American Mercury, Gemini, and Apollo programs. There were no fatalities attributed to this landing method although a Mercury astronaut almost drowned.

The major disadvantage of water landing is the significantly increased cost and time to refurbish a RLV for another flight and the cost of a recovery ship. A large number of people would be required to recover the vehicle and landing in water would mean that the vehicle would need to be carefully disassembled and repaired after every flight. Currently there is no reusable thermal protection system that can withstand a dunking into the water.

Another disadvantage is that some tourists may become sea sick after their flight.

Retrorockets. The Russians have successfully used retrorockets to cushion the final touchdown impact for all their manned space flight missions. On the Soyuz capsule, the retro rockets are placed behind the heat shield to protect them from reentry heat. The heat shield is dropped at altitude, which not only exposes the retro rockets but also prevents the hot heat shield from heating up the cabin interior. The heat shield becomes a hazardous piece of debris.

For optimum weight, the rate of descent for a parachute retrorocket system is in the order of 35 fps (10 m/s) or above. This rate of descent is too high for manned vehicles if the requirement exists for minimum aircrew injury at retrorocket malfunction. Hence retrorockets must be very reliable. The retrorocket component weight for a 35 fps (10 m/s) parachute descent is approximately 2% of the overall landing weight.

Finally there is a possibility of a ground fire caused by the retrorockets.

Airbags. Kistler Aerospace is planning on using airbags to cushion the parachute landing of their unmanned K-1 RLV. Kistler sized its parachutes for a 20 fps (6 m/s) rate of descent. The only manned operational use of parachutes and airbags was the F-111 crew escape capsule. This system had a very high landing impact at 26 fps (8 m/s) vertical, which was equal to a 10.5 ft (3 m) fall. As a result, a high percentage of crew were injured (of those that survived) during an ejection. The F-111 airbag system weighed approximately 3% of the escape capsule's weight. This system was also considered for the B-1 bomber, but standard ejection seats were chosen instead.

Crushable Impact Attenuators. Crushables include balsa wood, several types of foams, and paper, plastic, and aluminum honeycomb. Balsa wood has the best energy absorption per pound of material weight (~24,000 ft-lbs per lb)12, followed by honeycombs, with foams having about 1/5 the capability of balsa wood. The rise of the force, called the onset rate in g per second, is important for manned vehicles, since the human body limits the onset rate. The Apollo moon lander, the LEM, used a crushable in its landing gear. Also the landing peak force (in excess of 10 g's) may determine the vehicle structural design, since landing impacts can be expected to higher than any other load a RLV will experience. To limit deceleration to 10 g's, the crushable must compress at least 30 inches (for a 28 fps parachute descent rate).

Pneumatic Retractor. The US Army is experimenting with a pneumatic retractor that would pull a parachute and its load together just before landing to reduce landing impact. The system is in its early development stage.

Parafoils

A parafoil is a rectangular ram-air lifting parachute. Its usage is limited to vehicles that have a landing weight less than 25,000 lbs (11,400 kg). The recently cancelled space station rescue vehicle, the X-38, used this landing method. Timing the landing flare is the main challenge. The X-38's average vertical landing velocity during 26 test landings was 20 fps (6 m/s), which is equal to a 6.2 ft (1.9 m) fall.[13] Its average during the last 4 test drops, which were considered by the test team to be very good, was 15 fps (4.6 m/s) vertical with an average impact acceleration of 12 g's. The hardest landing was at 27 fps (8.2 m/s) and 41 g's. These are hard landings. As a comparison, naval aircraft hit an aircraft carrier flight deck at 10 fps (3 m/s) vertical. The X-38 landing gear was damaged or destroyed often during its test landings.

A parafoil's landing rate of descent is not much different from that available from much lighter circular parachutes. Parafoils are heavier than circular parachutes because they have more than twice the fabric (for upper and lower surfaces plus vertical ribs). The X-38's parafoil alone weighed 9.8% of the X-38's landing weight as compared to the Apollo parachute's 2.8% of capsule weight. In addition to the parafoil, the weight of the landing gear, drogue chutes, and parachute compartments must be accounted for when comparing parafoils to other landing modes.

Rigid or semi-rigid decelerator.

German DaimlerChrysler and Moscow-based NPO Lavochkin have developed a cone shaped aerodynamic decelerator known as IRDT (Inflatable Reentry Descent Technology). It was reentry flight tested in 2000 with an experimental Russian booster rocket and it survived its return to Earth, although it was damaged on the way down. IRDT technology is not intended for manned flight. Such a decelerator weighs as much as a wing and wheel system and would require either an airbag or retrorocket to cushion its final impact. Probability of falling over is very high due to the large amount of area presented to ground winds.

Rocket Recovery

The McDonnell Douglas DC-X is an example of a rocket powered vertical lander and tailsitter. In the DC-X, large amounts of propellant had to be reserved for landing which limited demonstrated peak altitude to only 8,700 ft (2,650 m). Furthermore during its test flights, the DC-X used its rocket engines as its primary control system, which meant that they could not be shut off. The DC-X was flown over desolate terrain because the rockets could ignite a fire in the ground foliage.

The next figure shows the relationship for propellant mass fraction, δp, and landing deceleration for a vertical lander. The figure was prepared assuming an Isp of 225 seconds and allowing no reserve propellant for hovering. The smallest possible δp required for landing is 0.07 (ballistic coefficient of 80 psf and deceleration of 1 g). The resulting maneuver is extremely aggressive and unrealistic - the rocket motors are not throttled up until the vehicle is 1,000 ft (300 m) above the ground and the motors are fired for only 8 seconds as the vehicle slows from 260 fps (80 m/s or 155 knots) to a hover. If the motors fail to throttle up, the vehicle crashes into the ground only 4 seconds later. As a comparison, an aggressive helicopter deceleration is about 0.15 g's. Realistic δp for rocket landings are likely to be 0.2 to 0.3 unless high Isp turbo pump fed engines are used. In other words, 20 to 30% of vehicle landing weight must be propellant.

The propellant mass fraction for a parachute retrorocket system is indicated by the 1 psf line (a vehicle descending under a standard Army cargo parachute). Notice less than 2% of the landing weight must be propellant for the retrorocket.

Rotor Recovery

There are two types of rotor recovery systems. Those that use stored rotor inertia only to cushion the landing and those that use tip rockets. A tip rocket powered rotor is more complex but would be lighter since a pound of tip rocket fuel has 20 times more energy than the stored kinetic energy in a pound of rotor tip weight.[15]

The Rotary Rocket Company experimented with a tip rocket powered rotor recovery system. They built a 63 ft (19 m) tall Atmospheric Test Vehicle (ATV) that flew 3 times in 1999 and demonstrated hover and forward flight with flight speeds up to 50 knots (84 fps or 26 m/s) and altitudes of 75 ft (23 m). The vehicle was found to be very unstable and extremely difficult to fly.[16] This landing method was abandoned when analysis showed that the rotor would have insufficient aerodynamic control power to control the vehicle through the high subsonic and transonic flight regime.

CANDIDATE ARCHITECTURES

Our study shows that vertical takeoff and some air launch methods are viable means of attaining the X-Prize altitude goal of 328,00 ft (100 km). Also wings with wheels and aerodynamic decelerators are technologically mature enough to serve as viable methods for recovery. As a result there are 4 architectures that can successfully win the X-Prize:

  1. Vertical takeoff, Aerodynamic decelerator
  2. Vertical takeoff, Wings with wheel landing
  3. Some Air Launch, Aerodynamic decelerator
  4. Some Air Launch, Wings with wheel landing

The preferred architecture for a vehicle that can not only win the X-Prize but also serve as a commercially successful sub-orbital tourist vehicle depends on the additional factors of safety, customer acceptance, and affordability.

SAFETY

Safety is the reasonable degree of freedom from those conditions that can cause injury, death to personnel, damage or loss of equipment or property; it is the freedom from danger. The advocacy of horizontal takeoff for many X-Prize concepts may be largely based on the intuitive belief that such vehicles would bring the greater safety of commercial airline aircraft to space transportation. In reality, direction of takeoff has little to do with safety. Previously built rocket powered airplanes do not appear to demonstrate remarkably better safety records as compared to vertical launch vehicles.

Sufficient data is available from expendable launch vehicles (ELVs) experience to pinpoint catastrophic failure modes sources for rocketpowered vehicles. The catastrophic failures during 1,176 ELV launches break down as:[17]


Sub system Failure Rate

Propulsion 69.7%
Avionics 16.3%
Staging 7.0%
Environmental 4.7%
Structural 2.3%

Propulsion Safety

Propulsion failures accounted for almost 70% of all the catastrophic failures. Elimination of propulsion failures requires picking the right rocket engine type. There are three major types of rocket engines. Liquid propellant engines use a separate oxidizer, for example, liquid oxygen; and a separate fuel, for example, kerosene or liquid hydrogen. Solid propellant motors use a solid propellant grain that contains both the oxidizer and the fuel. Finally a hybrid motor typically uses a liquid oxidizer such as nitrous oxide or liquid oxygen and a separate solid fuel grain such as rubber, wax, or plastic.

Liquid propellant engines can fail either catastrophically or benignly. Roughly 3 out of 4 kerosene and liquid oxygen ( LOX) engine failures are benign, in the sense that the engine failure results in loss of thrust, but not in immediate destruction of the vehicle. The historical benign failure ratio for kerosene- LOX engines is 0.6% (benign failures / engine flights) while the catastrophic failure ratio is 0.2%. While the number of solid rocket motor failures is small, they have all been catastrophic, rather than benign failures.[18] Hence designers should limit the number of liquid engines or solid motors in a vehicle since a large number of either would increase the chance of a catastrophic failure.

On the other hand, according to the Department of Defense Explosives Safety Board, hybrid motors can be fabricated, stored, and operated without any possibility of explosion or detonation.[19] Other advantages include the ability to be stopped, restarted, and throttled; easy (and hence potentially cheaper) ground handling; and relative insusceptibility to grain flaws.

Multiple hybrid motors can be used to improve propulsion system reliability by providing redundancy in the case that one hybrid motor fails to produce adequate thrust. Such active redundancy (also referred as to as 'engine out' and 'fail-safe' capability) can allow a vehicle to complete a flight following the failure of any one engine, provided of course that the engine failure is benign. In this respect, hybrid motors are like jet engines, which have an almost zero probability of catastrophic failure, but sometimes may not produce thrust. Note that hybrid motors and jet engines can catch on fire, so a fire suppression system is needed.

Experience has also shown that a significant number of propulsion failures occur during engine start. Although there has been some good progress in understanding steady state combustion process, the prediction and modeling of a new rocket engine during a start up, especially those using turbo pumps, is immature.

Historically, air launched RLV engines are started after the RLV is dropped from the carrier aircraft to ensure carrier aircraft safety. If the engine does not start or has some other malfunction that would require an immediate engine shutdown then propellant must be dumped and an emergency landing must be completed.

On 9 November 1962, Jack McKay was seriously injured and the X-15 severely damaged when its engine failed to throttle up after release from the carrier B-52 aircraft. The X-15's landing gear failed during the emergency landing due to overstress caused by a faster than normal landing speed (296 mph instead of the normal 230 mph) which was caused by an incomplete propellant jettison.[11]

Experience has shown that the initial few seconds after ignition tends to determine whether a rocket engine will or will not fail for the remainder of its operation. This means that in the case of a ground launch, a rational way to prevent vehicle failure would be to hold the vehicle on the pad until all systems appear to be in order. In event of problem, a controlled shutdown on the pad can be executed. Pad hold-down, generally for 5 seconds after ignition, makes it possible in the event of engine failure to shut down all engines and abort the flight on the pad.

Once released from the takeoff pad, the engines on a vertical takeoff vehicle must operate until the vehicle is high enough to dump its remaining propellant and conduct an emergency landing -- an altitude that is likely to be at least 10,000 ft (3,000 m) above the ground. Using active redundancy with multiple hybrid motors and sizing the motor's thrust so that the vehicle can fly away with a motor out can allow a vertical takeoff vehicle to complete an abort as long as a single motor failure does not cascade into other failures. For example, a vehicle with 4 hybrid motors and a lift-off thrust-to-weight ratio (T/W) of 1.6 can climb away in the event of a motor failure, even at lift-off.

Advocacy of horizontal takeoff may be based on the mistaken notion that such a vehicle is safer at takeoff. However, propulsion failure at takeoff rotation will also most likely result in a loss of vehicle (for example, the Sanders-Roe SR-53 takeoff fatality). There is no data that supports the contention that horizontal takeoff reduces rocket engine propulsion failure causes or rates. In addition, horizontal takeoff adds failure modes not found in the other takeoff methods. For example, 4 of the 19 SR-71 Blackbirds destroyed in accidents were due to blown tires on takeoff.

Avionic Safety

Aircraft practices can be used to minimize avionic failure modes in all four viable sub-orbital architectures. Installing redundant systems can reduce avionics and electrical failures, as this is done in commercial aircraft. Most ELV's have only single string avionics and electrical systems. In ELV's, redundancy to ensure mission success has been relegated to duplication of the complete ELV.

Staging Safety

The X-Prize altitude goal of 328,000 ft (100 km) can be reached with a single stage vertical takeoff vehicle. Air launch introduces the additional possibility of a separation failure, such as a RLV and carrier aircraft collision. Hence vertical takeoff is favored over air launch to improve staging safety.

In some concepts that use parachute or parafoil recovery, the crew compartment separates from the rest of the vehicle to soften the landing impact for the crew. Hence crew compartment separation represents a critical failure mode -- one that is difficult to minimize with a redundant design. Also the historical fatality rate for sport parachute jumping is much higher than that for landing an airplane or glider. Hence a winged landing is favored over an aerodynamic decelerator to improve staging safety.

Environmental Safety

Launching in bad weather causes environmental related failures; for example, the Challenger shuttle accident. Tourist vehicles do not have to be launched in bad weather and they do not have to make an orbital launch window or time slot. Again wings are favored over aerodynamic decelerators since the wind limit for landing with wings is much higher than that for landing with either a parachute or parafoil. Parachute landing into the water adds the consideration of sea state; large waves could sink the RLV or drown the passengers.

Structural Safety

ELV's are built with a small structural factor of safety (typically 1.2) because they have to be light enough to make it to orbit. A sub-orbital vehicle can afford more empty weight; hence structural factors of safety can be increased to commercial aircraft standards (typically 1.5), which have been proven to reduce structural failures.

CUSTOMER ACCEPTANCE

Prospective customers are expected to be spectators first and will observe several takeoffs and landings before they purchase a ride. They would also like their family and friends to watch their flight. One reason for the numerous horizontal takeoff X-Prize entrants may be the perception that customers would prefer horizontal takeoff. Unfortunately such vehicles cannot reach X-Prize altitudes with today's technology.

Vertical launch provides a spectacular event that can serve as the centerpiece of a space based theme park or an air show. It can be easily filmed and televised. In contrast, air launching at 20,000 to 50,000 ft (6 km to 15 km) may provide very little for a spectator located on the ground to watch.

Similarly a winged landing along a steep glide path onto a runway located adjacent to the takeoff pad some 15 to 20 minutes after the launch is another exciting event. A parachute landing, perhaps miles away from the launch site or into the water may hold little interest for a spectator.

AFFORDABILITY

The costs to operate a sub-orbital tourist vehicle are highly dependent on launch rate. The more an RLV flies, the cheaper it gets. This may create a chicken and egg dilemma. The investment to build a vehicle cannot be justified unless there is a demand for many flights, but the demand for flights depends on a low ticket price.

The ticket price for a ride on a sub-orbital tourist vehicle also depends on the number of passengers carried per flight since fixed costs such as the pilot and the mechanic's salary can be divided among a larger number of ticket fares as the vehicle becomes larger. Also vehicles do not scale directly with number of passengers -- a 4 passenger vehicle weighs less than twice as much as a 2 passenger vehicle. Hence, vertical launch and winged landing are favored since there is no limit to the size of the vehicle using these takeoff and landing modes.

Affordability is also driven by the overall complexity of the vehicle systems, which in turn determine operational support requirements. Advocacy of horizontal takeoff for many X-Prize concepts may have been based on the intuitive belief that such vehicles would bring the improved operational features of airplanes to space transportation. In reality, vertical takeoff concepts are actually less complex than either horizontal takeoff or air launch because of fewer systems, i.e. no jet engines (in addition to the rocket engines), no landing gear to retract, no carrier aircraft or release mechanisms (for air borne launch), and no multiple propellant tanks and transfer pumps to manage center of gravity location.

When comparing landing methods, a vehicle using winged landing would take less effort to turn around for the next flight as compared to a vehicle using a parachute or parafoil recovery. The X-38 parafoil weighed almost 2,000 lbs (900 kg) and took weeks to inspect and repack. Because parachutes and parafoils have a limited weight capability, the RLV may also have to separate into several parts, which would further increase the effort necessary to ready it for another flight.

Thus for the four viable architectures capable of the X-Prize, vertical takeoff with winged landing should take the least effort to turn around for the next flight. Air launching with parachute landing into the water should take the most effort to get ready for another flight. The air launched RLV also requires a custom carrier aircraft that must be developed and operated and it requires two pilots, one for the carrier aircraft and another for the RLV. For commercial sub-orbital flights, FAA certification costs will be double since both the carrier aircraft and RLV must be certified.

However, costs also depend on the size of the vehicle. Here the order is reversed. An air launched RLV can be the smallest RLV with a launch weight between than 6,000 lbs to 7,000 lbs (2,700 kg to 3,200 kg) for a 3 person craft, not including the weight of the carrier aircraft. The heaviest architecture is vertical takeoff with winged landing with a takeoff weight in the order of 40,000 lbs to 50,000 lbs (18,000 kg to 23,000 kg). The other architectures have takeoff weights that are between these weights. Obviously propellant cost is a function of takeoff weight, thus air launching would have the lowest propellant cost.

Thus there is no clear resolution on which architecture would have the lowest cost to develop and operate.

CONCLUDING REMARKS

This paper shows that both vertical takeoff and some air launch methods are viable means of attaining sub-orbital altitudes and wings and aerodynamic decelerators are viable methods for recovery. These conclusions are based on statistical methods using historical data coupled with time-stepped integration of the trajectory equations of motion. As a result there are 4 architectures that can successfully win the X-Prize.

Based on the additional factors of safety, customer acceptance, and affordability, we also feel that the preferred architecture for a commercial vehicle that will successfully and profitability carry tourists on sub-orbital flights is an architecture that uses Vertical Takeoff and winged un-powered Horizontal Landing onto a runway ( VTHL) powered by hybrid rocket motor propulsion.

This preferred configuration would have the following features:

Vertical Takeoff -- Vertical takeoff provides the capability to reach X-Prize and commercially viable altitudes. In addition, it has significant growth capability to reach higher altitudes and carry large number of passengers. Vertical takeoff provides a spectacular launch event for spectators.

Single Stage -- Single stage can be used for sub-orbital flights. Single stage eliminates parts falling off the vehicle that would restrict operation to either a government range or over the ocean flight. Finally using a single stage architecture eliminates staging separation catastrophic failure modes.

Multiple hybrid motors -- This propulsion arrangement can eliminate catastrophic propulsion failure modes when coupled with a takeoff pad hold-down system that allows motor health to be determined during the first few seconds of motor operation. In event of a problem, the motors can be shutdown. In this way, vehicle reliability can be greatly enhanced.

Wings and Wheeled Landing -- This landing mode can provide an airliner-like soft landing at an intended point and provide a spectacular recovery event for spectators.

Rockwell's 1995 X-33 Concept (NASA image)

We expect that a commercially successful suborbital tourist vehicle would look like a scaled down version of Rockwell's 1995 vertical takeoff and horizontal landing concept for the X-33 program. However, our concept would use the outer mold line of the later NASA X-34 because of the quality of public domain wind tunnel and computational fluid dynamic (CFD) data published.[20] Producing this data cost NASA approximately $16 million and it will significantly reduce the cost and risk of developing a space tourist vehicle.

During the past year, we have been working on a hybrid rocket powered VTHL concept that uses the X-34's outer mold line. We have completed a conceptual design on all the vehicle's subsystems, trajectory analysis, weight and balance, and layout drawings. A description of our hybrid powered VTHL can be found in our patent application and a 150 page PowerPoint presentation provides details of our concept.[21] We can provide briefings to parties that would be interested in helping us establish sub-orbital tourism. Our current efforts include students working on a high-fidelity flight simulation and on a detailed finite element model of the vehicle's structural design.

X-34 Outer Mold Line (NASA image)

We believe that a successful sub-orbital tourist vehicle will take the cooperation of industry, federal and state governments, and universities. Governments will provide regulatory stability, facilities such as airports, NASA data, tax credits, jobs, and a favorable environment for business. Universities will provide "out-of-box" thinking, research, and future employees. Innovative companies will actually build and market the vehicle, create new jobs, products, services, profits, and pay taxes that ultimately fund the other two sectors.

Finally, establishing space tourism as a commercial business will require taking some business risks. However, we hope that this paper shows how to avoid taking unnecessary technical risks by selecting a vehicle architecture that works and avoiding those architectures that violate physical laws or are beyond current state of the art.

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M Sarigul-Klijn & N Sarigul-Klijn, 2003, "Flight Mechanics of Manned Sub-Orbital Reusable Launch Vehicles with Recommendations for Launch and Recovery", AIAA 2003-0909. January 2003 (Revised April 2003).
Also downloadable from http://www.spacefuture.com/archive/flight mechanics of manned suborbital reusable launch vehicles with recommendations for launch and recovery.shtml

 Bibliographic Index
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