The following is an essay published in The Space Review on March 31, 2008. This essay responds to a earlier Space Review essay by Eric Hedman.
Assessing the Practicality of Scramjet-Powered, Single-Stage Aerospaceplanes
By Mike Snead
Clearly, the United States needs improved space access with “aircraft-like” safety and operability. In his Space Review article, “Space launch evolution and revolution” (The Space Review, December 10, 2007), Eric Hedman advocates focusing American space access development efforts on scramjet-powered, fully-reusable aerospaceplanes. He states: “The Holy Grail of low per passenger cost launcher concepts keeps coming back to scramjets.”
Scramjet-powered space access was first investigated, in the United States, in the late 1950’s and early 1960’s as part of the Department of Defense’s first aerospaceplane studies. These studies started what has now been nearly a half century of research into the design and performance modeling of scramjet engines and scramjet-powered flight systems.
From a top-level system design and operability perspective, a single-stage aerospaceplane is certainly, as Hedman highlights, the Holy Grail. As the United States struggles in its transformation into a true spacefaring nation, there is a critical need to identify a practical strategy to meet the need for aircraft-like, fully-reusable space access. Should scramjet-powered aerospaceplanes be placed front and center in the pursuit of the needed improved space access capabilities or is another approach the more practical approach at this time? As with any quest, success lies with selecting the more readily implementable path. This article summarizes some of the challenges involved in the scramjet-powered aerospaceplane path.
All airbreathing jet engines have four primary segments – inlet, compressor, combustor, and nozzle. What is unique about the scramjet is that the airflow velocity through the combustor is designed to exceed the local speed of sound. “Scramjet” is an abbreviation for “supersonic combustion ramjet.”
In turbojets and ramjets, the velocity in the combustor is kept below the local speed of sound to enable efficient fuel injection, mixing, and combustion. As the flight velocity increases, useful energy in the incoming air is lost—converted to excessive heat that must be actively cooled—as the air flow is decelerated in the inlet and compressor to local subsonic speeds for entry into the combustor. This subsonic combustor speed limit effectively places an upper bound on aircraft’s flight velocities of around Mach 3-5 for pure turbojets and around Mach 6-7 for pure ramjets.
The scramjet reduces this loss by intentionally keeping the local flow speeds through the engine supersonic. If the scramjet engine and vehicle can be designed to withstand the flight external and internal temperatures and pressures, and if fuel combustion can be maintained, then very high airbreathing flights speeds, perhaps as high as Mach 15, may be achievable.
Research efforts to advance basic scramjet propulsion technologies, first begun in the 1960’s, continue. In 2004, for example, the 1,400 kg, 4.7 m-long, scramjet-powered, X-43A achieved an airbreathing speed of Mach 9.6 for about 10 seconds. A large rocket was used to boost the X-43A, following its release from the B-52 carrier aircraft, to the hypersonic test conditions needed to perform the scramjet engine’s tests.
Space access design considerations
For typical space launch systems using liquid-fueled rocket engines, both the fuel and oxidizer must be carried internally. The Space Shuttle does this with its large External Tank. Through the addition of airbreathing propulsion, some of the oxidizer is instead provided by the atmosphere. Ideally, the use of airbreathing as well as rocket propulsion enables the mass of the oxidizer and oxidizer tanks to be reduced. This approach, it is anticipated, will yield an aerospaceplane design that has a reduced takeoff gross weight when compared with the pure rocket baseline design.
The apparent advantage of using scramjets for aerospaceplanes is that the maximum flight velocity, where air still provides the needed oxidizer, can be high. Theoretically, a significant reduction in the propellant mass fraction can be achieved. The airbreathing “Holy Grail” Hedman notes is that such a reduction may enable a “closed design” for a single-stage aerospaceplane to be achieved. As described in the following, this path has many challenges that have yet to be successfully addressed. It is worth noting that closed designs of pure rocket-powered, single-stage aerospaceplanes, capable of achieving aircraft-like safety and operability using current technologies, have also not been proposed.
Single-stage space access system mass fractions
To better understand the design challenges of achieving a closed single-stage aerospaceplane design, an understanding of the concept of mass fractions is needed.
A single-stage aerospaceplane’s takeoff mass can be divided into two categories—propellant mass fraction and empty mass fraction—usually expressed as a percentage of the takeoff mass. The sum of the two fractions, by definition, is 100 percent. As one increases, the other must decrease.
As the name implies, the propellant mass fraction is the percentage of the aerospaceplane’s takeoff mass that is propellant—fuel and oxidizer. The propellant mass fraction required to achieve orbit is mathematically calculated from the change in velocity needed to achieve orbit, the losses due to drag and trajectory that must be accounted for, and the propulsion subsystem’s net efficiency in converting the fuel’s energy into increasing vehicle velocity.
The net propulsion efficiency is the net thrust produced per unit mass of propellant consumed. This is the familiar “specific impulse” term that is used, along with the thrust, to describe the performance of a rocket engine. It can also be used for airbreathing engines. By using oxygen from the air, instead of from onboard tanks, airbreathing engines ideally increase the net propulsion efficiency by decreasing the mass of propellant consumed each second to yield the same thrust. This has the benefit of reducing the theoretical propellant mass fraction.
For a rocket-powered, single-stage aerospaceplane, the theoretical propellant fraction to achieve orbit will range from about 88-95 percent. This means that 88-95 kg of every 100 kg of takeoff mass must be propellant if orbit is to be achieved. Rocket-powered aerospaceplanes are essentially large flying propellant tanks.
Ideally, when a very efficient airbreathing-rocket engine combination is used, the theoretical propellant fraction reduces to about 60-70 percent. As a result, the empty mass fraction for the scramjet solution (30-40 percent) can be two to eight times larger than that for a pure rocket design (5-12 percent). From a design perspective, this provides a substantial apparent advantage to the airbreathing approach. It seems obvious that an airbreathing design where 30-40 kg of every 100 kg of takeoff mass can be hardware would be easier to design and build compared with a pure rocket design that only permits 5-12 kg of every 100 kg of takeoff mass to be hardware.
By subtracting the theoretical propellant fraction from 100 percent, the remaining fraction is the empty mass fraction objective—what the design engineers need to achieve. If a safe, operable, and buildable design can be defined where the predicted empty mass fraction is equal to or less than this objective, the design is said to “close.”
Challenges in closing the design of single-stage, scramjet-powered systems
Today, the American aerospace industry has the ability to design and build two-stage, fully-reusable space access systems that are believed to “close” based on the results of government and industry conceptual design studies using current technologies. Generally, these designs use a vertical takeoff, horizontal landing approach that is pure rocket powered. The original fully-reusable design for the Space Shuttle from the early 1970’s, before the design was changed to the expendable External Tank and Solid Rocket Boosters, is an example of this configuration.
While various ideas for single-stage, scramjet-powered aerospaceplanes have been proposed since the 1960’s, closed single-stage designs using current technologies and capable of aircraft-like safety and operability have not yet been achieved. Why is this? The problem is that the use of scramjets adds significant complexity to the design and verification of a single-stage aerospaceplane while at the same time adding uncertainty to the estimation of the required propellant fraction. Essentially, it makes the problem of predicting design closure and then demonstrating a successful design much harder, as discussed in the following examples.
– Added propellant tank mass
One design challenge is that scramjet airbreathing propulsion adds significant additional empty mass that is not needed for a pure-rocket system. While the addition of airbreathing propulsion may decrease the required propellant fraction, the average propellant density may also decrease. This can result in disproportionally larger total propellant tank volume and larger tank total mass.
This design dilemma results from the fact that a scramjet requires hydrogen as the fuel for flight speeds greater than about Mach 7. Hydrogen, it is hoped, will have the ability to sustain combustion in the scramjet’s supersonic combustor at these higher Mach numbers where no other fuel appears to be able to sustain combustion unaided. Liquid hydrogen, however, has the disadvantage of having a very low density—a beverage cup that holds 16 oz. of water only holds about 1 oz. of liquid hydrogen. Liquid hydrogen also is very cold with a temperature close to absolute zero. This requires added tank insulation and special fuel tank pressurization management. Increased tank volume, increased tank insulation, and special fuel tank pressurization management generally yields increased tank mass per kg of propellant carried.
– Added thermal protection subsystem mass
Closely related to the likely increase in total propellant tank mass is the additional thermal protection subsystem (TPS) mass required for hypersonic scramjet operations. All reusable space access systems require thermal protection during reentry. During ascent under rocket power, the thermal loads are comparatively moderate and no additional thermal protection is normally required. The thermal protection required during scramjet-powered ascent can, however, exceed that required for reentry. As a result, the total TPS mass will usually increase, compared with a pure-rocket design, due to the higher temperatures, higher aerothermal loads, the increased need for active cooling, and the larger total propellant tank surface area.
– Added propulsion subsystem mass
Another empty mass design issue is the increase in the mass fraction of the propulsion system to incorporate airbreathing propulsion. In oxygen-kerosene rocket engines, for example, the installed engine thrust-to-weight ratio is about 60-80 while it is about 50-60 for oxygen-hydrogen rocket engines. The installed thrust-to-weight of the best supersonic turbofans is about 5-8 and about 8-12 for the best subsonic turbofans. The complexity of the design of the scramjet, combined with the need to retain a rocket engine capability, means that the installed thrust-to-weight of the integrated airbreathing-rocket propulsion subsystem for an aerospaceplane may only be in the range of 3-5. Compared with a pure-rocket design where the engines may comprise about 2-3 percent of the takeoff mass, the integrated airbreathing-rocket engine may 10X heavier, comprising 20 percent or more of the takeoff mass, or about half of the allowable empty mass fraction.
– Uncertainty in predicting the required propellant fraction
Design challenges also arise on the propellant fraction side of the design closure ledger. To develop a thorough closed design, good confidence in the accuracy of the required propellant fraction is needed—typically, 3-4 significant figures. This then enables the design engineers to confidently calculate the empty mass fraction objective and have a good definition of the specific flight conditions that need to be addressed—dynamic pressures, aerothermal loads, etc.
Unfortunately, modeling and verifying the performance of an integrated airbreathing propulsion subsystem from Mach 0-15 remains very challenging. Limited ground and flight scramjet test data means that there remains substantial uncertainty about the actual achievable performance, especially at high Mach numbers. While further scramjet technology development and testing may resolve this issue, substantial further improvement in scramjet performance modeling is required before the conceptual design closure of scramjet-powered aerospaceplanes can be achieved.
Scramjet testing design and verification challenges
Rocket and airbreathing engines are extensively tested to verify performance, safety, and operability. New jet engine designs routinely undergo thousands of hours of ground test while 500-1000 test firings of a new rocket engine design is the norm. For an airbreathing propulsion system that includes scramjets, the wide range of airbreathing operation from Mach 0 to 15 will require more extensive testing than is typically done for jet engines.
One challenge hindering scramjet development is that ground test facilities for full-scale testing of airbreathing propulsion systems are quite limited for Mach numbers greater than about Mach 3-4. Achieving the correct combinations of airflow velocity, mass flow rate, temperature, and sustained test conditions in test sections of the needed size for large scramjets is very difficult. Because of this, more extensive use of flight testing would be the presumed alternative approach. As the recent X-43A test program has shown, scramjet flight testing is very expensive and time consuming. The seven-year X-43A program, using three small expendable test vehicles, cost approximately $230 million. Two of the three tests conducted were successful in demonstrating scramjet operations for a total of about 30 seconds of test data. Relying on flight testing to very high Mach numbers, assuming this is successful, will add significantly to the length of time and cost required to successfully develop a scramjet for an aerospaceplane.
Scramjet-influenced airframe structural design and verification challenges
A key element of a successful airframe structural design is one that can be economically built and tested. Fabrication and test considerations are as important to the airframe’s design as are the aero-thermal-inertia loads. An innovative structural design engineer may be able to conceptualize a highly efficient, low-mass airframe design, but it may not be able to be successfully built and/or tested. This is especially important for aerospaceplanes because of the significant aerothermal loads acting on the vehicle during descent and, with scramjet propulsion, during ascent.
An example of the design and verification testing challenges that will arise with scramjet-powered aerospaceplanes comes from the National Aerospace Plane (NASP/X-30) program of the late 1980’s. It focused on a scramjet-powered, horizontal takeoff and landing, single-stage system that would have a takeoff mass of about a half-million kilograms—about the same as the new Airbus A380.
New airframe designs go through a series of component and full-scale ground tests to verify structural integrity. This is done for X- and Y-aircraft’s airframe design as well as for the production aircraft’s airframe design. Such testing involves the initial static tests to establish the strength and stiffness of the structure followed by cyclic testing to establish the damage tolerance and durability of the structure. The need for scramjet-powered systems to carry liquid hydrogen, combined with the time-dependent aero-thermal loads from the use of high-Mach airbreathing propulsion during ascent, adds significant complexity to the airframe’s structural testing.
While most airframe structural testing is conducted in ambient air environments, this will probably not be possible for scramjet-powered aerospaceplanes. Because no other liquid properly simulates the thermal response and mass characteristics of the liquid hydrogen in the fuel tank, structural testing of the airframe of scramjet-powered aerospaceplanes may be expected to require testing with, perhaps, several hundred tons of liquid hydrogen onboard. Any type of structural testing involving flammable liquids poses additional safety constraints. Testing of hydrogen is, perhaps, the most difficult because of hydrogen’s flammability in air at low concentrations and hydrogen’s ability to escape through small leaks when other gases cannot—such as small cracks in the tank walls or internal piping. The risk of even small structural failures that release hydrogen requires that such testing be done in special inerted facilities, possible the size of a domed football stadium, located several kilometers away from inhabited buildings.
High Mach airbreathing propulsion also introduces the variable of time in the structural verification testing because the internal airframe/tank temperatures and internal loads are time dependent. When combined with the changing external aero-thermal-inertia loads acting on the airframe as the vehicle accelerates to the Mach 15 scramjet cutoff velocity, this presents a very complex structural test matrix consisting of, potentially, hundreds of different test points covering both ascent, reentry, and emergency flight conditions.
Using past plans for NASP/X-30 structural testing as the example, the quantity of liquid hydrogen in the tanks would be set to the amount remaining at that point in the flight profile and the internal structure would be brought to the design internal temperature conditions using, probably, internal electrical heating elements. The external thermal and mechanical loads would then be applied over hundreds of square meters of surface area. At the same time, the airframe and scramjet’s active cooling system, using gaseous hydrogen at pressures up to 5,000 psi, would be started to simulate the cooling of much of this external surface area to ensure the correct airframe internal thermal loads and structural deformation. Obviously, such structural testing involving large quantities of combustible fluids near absolute zero, radiant thermal loads developing external TPS temperatures of over 1000 F, hot hydrogen gas at high pressures and temperatures flowing through hundreds of square meters of minimum-gage, actively cooled structures, and large mechanical loads that induce significant structural deflection, will be challenging to undertake.
Further complicating the structural integrity tests is that they will not be comprised of a single test for each design point. As mentioned, the initial static strength and stiffness verification of the structural integrity of the airframe may involve hundreds of test points to release the system for flight test and support envelope expansion. Some of these test points will probably need to be retested using actual flight-measured temperature and aero loads. Also, verification of the resistance of the airframe to loss of structural integrity brought about by temperature cycling and operational usage will involve further repetitive load testing of hundreds to thousands of cycles when the time comes to verify the structural integrity of the production configuration.
These structural design and testing challenges indicate that, as with scramjet engine testing, the structural development of scramjet-powered aerospaceplanes will be much more complex than that required for the more traditional airframe design expected to be used in two-stage, rocket-powered aerospaceplanes.
Scramjet-powered aerospaceplane program length and cost
The final challenge facing the development of scramjet-powered aerospaceplanes is the length and cost of their development programs. One estimate, prepared in 2003, stated that the projected cost through 2018 for “large-scale scramjets and turbojets,” as part of the National Hypersonics Initiative, was $8-14 billion. (Turbojets would be needed to accelerate the vehicle to the Mach 4-7 speed necessary for the ramjet/scramjet to start to operate.) The additional technology development demands for the airframe, thermal protection system, propellant tanks, and other flight-critical subsystems influenced by the use of hypersonic airbreathing propulsion will add significantly to the necessary technology risk reduction investments that would also be required.
A primary consideration in laying out a projected development timeline for a scramjet-powered aerospaceplane is the recognition that much of the needed technology development and maturation will involve substantial flight testing. The reality of scramjet propulsion is that test size counts due to the influence of the boundary layer characteristics on the engine operation. Hence, small test engine sizes, such as used on the X-43A, have inherent limits on their contribution to technology maturation. Near full-scale engines will be needed to verify engine performance and operability and near full-scale flight systems will be needed to demonstrate the ability to actually achieve orbit.
It may be expected that following the completion of the preliminary round of technology development efforts—perhaps by 2020 assuming sufficient funding—a near full-scale X-aircraft using all representative flight hardware designs, covering the full airbreathing range from Mach 0 to Mach 15, would be built. This would be followed by a pre-production Y-aircraft that is near full-scale and includes the rocket propulsion system to verify the ability to achieve orbit and reenter safely. Once the Y-aircraft program is successfully completed, the production program would start. With a typical design, build, test cycle of 10 years for advanced flight systems and starting in 2020, the X-aircraft program could be completed in 2030, the Y-aircraft program in 2040, and the production system could become operational about 2050.
With the impact of the upcoming termination of Space Shuttle operations on American space operations, it is very apparent that the United States needs substantially improved passenger and cargo space access. It is also clear that only a fully-reusable system design will enable improved “aircraft-like” safety and operability to be achieved. The alternative to the scramjet-powered, single-stage aerospaceplane approach is the two-stage, rocket-powered, fully-reusable aerospaceplane. The technologies required to start development of this approach are sufficiently mature that production development of the first generation could start now. By 2020, perhaps as early as 2016, such a two-stage aerospaceplane system could become operational and America’s ability to access space will leap forward more than three decades earlier than pursuing the scramjet-powered, single-stage approach.
Getting passengers and cargo to space with “aircraft-like” safety and operability are the primary needs of a true spacefaring nation. Hopefully, further technology development will yield single-stage capabilities in the future. While such technology development should continue, it is important that America’s future in space not be held hostage by the imposition of the scramjet technology path to space. Achieving aircraft-like safety and operability are the only design prerequisites and these will bring the desired lower costs. Starting the development of rocket-powered, two-stage, fully-reusable aerospaceplanes is, today, America’s readily implementable option to meet its space access needs as a true spacefaring nation.