Single-stage-to-orbit

From Wikipedia, the free encyclopedia

https://en.wikipedia.org/wiki/Single-stage-to-orbit

The VentureStar was a proposed SSTO spaceplane.

A single-stage-to-orbit (SSTO) vehicle reaches orbit from the surface of a body using only propellants and fluids and without expending tanks, engines, or other major hardware. The term usually, but not exclusively, refers to reusable vehicles. To date, no Earth-launched SSTO launch vehicles have ever been flown; orbital launches from Earth have been performed by either fully or partially expendable multi-stage rockets.

The main projected advantage of the SSTO concept is elimination of the hardware replacement inherent in expendable launch systems. However, the non-recurring costs associated with design, development, research and engineering (DDR&E) of reusable SSTO systems are much higher than expendable systems due to the substantial technical challenges of SSTO, assuming that those technical issues can in fact be solved. SSTO vehicles may also require a significantly higher degree of regular maintenance.

It is considered to be marginally possible to launch a single-stage-to-orbit chemically fueled spacecraft from Earth. The principal complicating factors for SSTO from Earth are: high orbital velocity of over 7,400 metres per second (27,000 km/h; 17,000 mph); the need to overcome Earth's gravity, especially in the early stages of flight; and flight within Earth's atmosphere, which limits speed in the early stages of flight due to drag, g, and influences engine performance.

Advances in rocketry in the 21st century have resulted in a substantial fall in the cost to launch a kilogram of payload to either low Earth orbit or the International Space Station, reducing the main projected advantage of the SSTO concept.

Notable single stage to orbit concepts include Skylon, which used the hybrid-cycle SABRE engine that can use oxygen from the atmosphere when it is at low altitude, and then using onboard liquid oxygen after switching to the closed cycle rocket engine at high altitude, the McDonnell Douglas DC-X, the Lockheed Martin X-33 and VentureStar which was intended to replace the Space Shuttle, and the Roton SSTO, which is a helicopter that can get to orbit. However, despite showing some promise, none of them has come close to achieving orbit yet due to problems with finding a sufficiently efficient propulsion system and discontinued development.

Single-stage-to-orbit is much easier to achieve on extraterrestrial bodies that have weaker gravitational fields and lower atmospheric pressure than Earth, such as the Moon and Mars, and has been achieved from the Moon by the Apollo program's Lunar Module, by several robotic spacecraft of the Soviet Luna program, and by China's Chang'e 5.

History

Early concepts

ROMBUS concept art

Before the second half of the twentieth century, very little research was conducted into space travel. During the 1960s, some of the first concept designs for this kind of craft began to emerge.

One of the earliest SSTO concepts was the expendable One stage Orbital Space Truck (OOST) proposed by Philip Bono, an engineer for Douglas Aircraft Company. A reusable version named ROOST was also proposed.

Another early SSTO concept was a reusable launch vehicle named NEXUS which was proposed by Krafft Arnold Ehricke in the early 1960s. It was one of the largest spacecraft ever conceptualized with a diameter of over 50 metres and the capability to lift up to 2000 short tons into Earth orbit, intended for missions to further out locations in the solar system such as Mars.

The North American Air Augmented VTOVL from 1963 was a similarly large craft which would have used ramjets to decrease the liftoff mass of the vehicle by removing the need for large amounts of liquid oxygen while traveling through the atmosphere.

From 1965, Robert Salkeld investigated various single stage to orbit winged spaceplane concepts. He proposed a vehicle which would burn hydrocarbon fuel while in the atmosphere and then switch to hydrogen fuel for increasing efficiency once in space.

Further examples of Bono's early concepts (prior to the 1990s) which were never constructed include:

• ROMBUS (Reusable Orbital Module, Booster, and Utility Shuttle), another design from Philip Bono. This was not technically single stage since it dropped some of its initial hydrogen tanks, but it came very close.
• Ithacus, an adapted ROMBUS concept which was designed to carry soldiers and military equipment to other continents via a sub-orbital trajectory.
• Pegasus, another adapted ROMBUS concept designed to carry passengers and payloads long distances in short amounts of time via space.
• Douglas SASSTO, a 1967 launch vehicle concept.
• Hyperion, yet another Philip Bono concept which used a sled to build up speed before liftoff to save on the amount of fuel which had to be lifted into the air.

Star-raker: In 1979 Rockwell International unveiled a concept for a 100-ton payload heavy-lift multicycle airbreather ramjet/cryogenic rocket engine, horizontal takeoff/horizontal landing single-stage-to-orbit spaceplane named Star-Raker, designed to launch heavy Space-based solar power satellites into a 300 nautical mile Earth orbit. Star-raker would have had 3 x LOX/LH2 rocket engines (based on the SSME) + 10 x turboramjets.

Around 1985 the NASP project was intended to launch a scramjet vehicle into orbit, but funding was stopped and the project cancelled. At around the same time, the HOTOL tried to use precooled jet engine technology, but failed to show significant advantages over rocket technology.

DC-X technology

Main article: McDonnell Douglas DC-X

The maiden flight of the DC-X

The DC-X, short for Delta Clipper Experimental, was an uncrewed one-third scale vertical takeoff and landing demonstrator for a proposed SSTO. It is one of only a few prototype SSTO vehicles ever built. Several other prototypes were intended, including the DC-X2 (a half-scale prototype) and the DC-Y, a full-scale vehicle which would be capable of single stage insertion into orbit. Neither of these were built, but the project was taken over by NASA in 1995, and they built the DC-XA, an upgraded one-third scale prototype. This vehicle was lost when it landed with only three of its four landing pads deployed, which caused it to tip over on its side and explode. The project has not been continued since.

Roton

Main article: Rotary Rocket

From 1999 to 2001 Rotary Rocket attempted to build a SSTO vehicle called the Roton. It received a large amount of media attention and a working sub-scale prototype was completed, but the design was largely impractical.

Approaches

There have been various approaches to SSTO, including pure rockets that are launched and land vertically, air-breathing scramjet-powered vehicles that are launched and land horizontally, nuclear-powered vehicles, and even jet-engine-powered vehicles that can fly into orbit and return landing like an airliner, completely intact.

For rocket-powered SSTO, the main challenge is achieving a high enough mass-ratio to carry sufficient propellant to achieve orbit, plus a meaningful payload weight. One possibility is to give the rocket an initial speed with a space gun, as planned in the Quicklaunch project.

For air-breathing SSTO, the main challenge is system complexity and associated research and development costs, material science, and construction techniques necessary for surviving sustained high-speed flight within the atmosphere, and achieving a high enough mass-ratio to carry sufficient propellant to achieve orbit, plus a meaningful payload weight. Air-breathing designs typically fly at supersonic or hypersonic speeds, and usually include a rocket engine for the final burn for orbit.

Whether rocket-powered or air-breathing, a reusable vehicle must be rugged enough to survive multiple round trips into space without adding excessive weight or maintenance. In addition a reusable vehicle must be able to reenter without damage, and land safely.

While single-stage rockets were once thought to be beyond reach, advances in materials technology and construction techniques have shown them to be possible. For example, calculations show that the Titan II first stage, launched on its own, would have a 25-to-1 ratio of fuel to vehicle hardware. It has a sufficiently efficient engine to achieve orbit, but without carrying much payload.

Dense versus hydrogen fuels

Hydrogen fuel might seem the obvious fuel for SSTO vehicles. When burned with oxygen, hydrogen gives the highest specific impulse of any commonly used fuel: around 450 seconds, compared with up to 350 seconds for kerosene.

Hydrogen has the following advantages:

• Hydrogen has nearly 30% higher specific impulse (about 450 seconds vs. 350 seconds) than most dense fuels.
• Hydrogen is an excellent coolant.
• The gross mass of hydrogen stages is lower than dense-fuelled stages for the same payload.
• Hydrogen is environmentally friendly.

However, hydrogen also has these disadvantages:

• Very low density (about 1⁄7 of the density of kerosene) – requiring a very large tank
• Deeply cryogenic – must be stored at very low temperatures and thus needs heavy insulation
• Escapes very easily from the smallest gap
• Wide combustible range – easily ignited and burns with a dangerously invisible flame
• Tends to condense oxygen which can cause flammability problems
• Has a large coefficient of expansion for even small heat leaks.

These issues can be dealt with, but at extra cost.

While kerosene tanks can be 1% of the weight of their contents, hydrogen tanks often must weigh 10% of their contents. This is because of both the low density and the additional insulation required to minimize boiloff (a problem which does not occur with kerosene and many other fuels). The low density of hydrogen further affects the design of the rest of the vehicle: pumps and pipework need to be much larger in order to pump the fuel to the engine. The result is the thrust/weight ratio of hydrogen-fueled engines is 30–50% lower than comparable engines using denser fuels.

This inefficiency indirectly affects gravity losses as well; the vehicle has to hold itself up on rocket power until it reaches orbit. The lower excess thrust of the hydrogen engines due to the lower thrust/weight ratio means that the vehicle must ascend more steeply, and so less thrust acts horizontally. Less horizontal thrust results in taking longer to reach orbit, and gravity losses are increased by at least 300 metres per second (1,100 km/h; 670 mph). While not appearing large, the mass ratio to delta-v curve is very steep to reach orbit in a single stage, and this makes a 10% difference to the mass ratio on top of the tankage and pump savings.

The overall effect is that there is surprisingly little difference in overall performance between SSTOs that use hydrogen and those that use denser fuels, except that hydrogen vehicles may be rather more expensive to develop and buy. Careful studies have shown that some dense fuels (for example liquid propane) exceed the performance of hydrogen fuel when used in an SSTO launch vehicle by 10% for the same dry weight.

In the 1960s Philip Bono investigated single-stage, VTVL tripropellant rockets, and showed that it could improve payload size by around 30%.

Operational experience with the DC-X experimental rocket has caused a number of SSTO advocates to reconsider hydrogen as a satisfactory fuel. The late Max Hunter, while employing hydrogen fuel in the DC-X, often said that he thought the first successful orbital SSTO would more likely be fueled by propane.

One engine for all altitudes

Some SSTO concepts use the same engine for all altitudes, which is a problem for traditional engines with a bell-shaped nozzle. Depending on the atmospheric pressure, different bell shapes are required. Engines designed to operate in a vacuum have large bells, allowing the exhaust gasses to expand to near vacuum pressures, thereby raising efficiency. Due to an effect known as Flow separation, using a vacuum bell in atmosphere would have disastrous consequences for the engine. Engines designed to fire in atmosphere therefore have to shorten the nozzle, only expanding the gasses to atmospheric pressure. The efficiency losses due to the smaller bell are usually mitigated via staging, as upper stage engines such as the Rocketdyne J-2 do not have to fire until atmospheric pressure is negligible, and can therefore use the larger bell.

One possible solution would be to use an aerospike engine, which can be effective in a wide range of ambient pressures. In fact, a linear aerospike engine was to be used in the X-33 design.

Other solutions involve using multiple engines and other altitude adapting designs such as double-mu bells or extensible bell sections.

Still, at very high altitudes, the extremely large engine bells tend to expand the exhaust gases down to near vacuum pressures. As a result, these engine bells are counterproductive due to their excess weight. Some SSTO concepts use very high pressure engines which permit high ratios to be used from ground level. This gives good performance, negating the need for more complex solutions.

Airbreathing SSTO

Skylon spaceplane

Some designs for SSTO attempt to use airbreathing jet engines that collect oxidizer and reaction mass from the atmosphere to reduce the take-off weight of the vehicle.

Some of the issues with this approach are:

• No known air breathing engine is capable of operating at orbital speed within the atmosphere (for example hydrogen fueled scramjets seem to have a top speed of about Mach 17). This means that rockets must be used for the final orbital insertion.
• Rocket thrust needs the orbital mass to be as small as possible to minimize propellant weight.
• The thrust-to-weight ratio of rockets that rely on on-board oxygen increases dramatically as fuel is expended, because the oxidizer fuel tank has about 1% of the mass as the oxidizer it carries, whereas air-breathing engines traditionally have a poor thrust/weight ratio which is relatively fixed during the air-breathing ascent.
• Very high speeds in the atmosphere necessitate very heavy thermal protection systems, which makes reaching orbit even harder.
• While at lower speeds, air-breathing engines are very efficient, but the efficiency (Isp) and thrust levels of air-breathing jet engines drop considerably at high speed (above Mach 5–10 depending on the engine) and begin to approach that of rocket engines or worse.
• Lift to drag ratios of vehicles at hypersonic speeds are poor, however the effective lift to drag ratios of rocket vehicles at high g is not dissimilar.

Thus with for example scramjet designs (e.g. X-43) the mass budgets do not seem to close for orbital launch.

Similar issues occur with single-stage vehicles attempting to carry conventional jet engines to orbit—the weight of the jet engines is not compensated sufficiently by the reduction in propellant.

On the other hand, LACE-like precooled airbreathing designs such as the Skylon spaceplane (and ATREX) which transition to rocket thrust at rather lower speeds (Mach 5.5) do seem to give, on paper at least, an improved orbital mass fraction over pure rockets (even multistage rockets) sufficiently to hold out the possibility of full reusability with better payload fraction.

It is important to note that mass fraction is an important concept in the engineering of a rocket. However, mass fraction may have little to do with the costs of a rocket, as the costs of fuel are very small when compared to the costs of the engineering program as a whole. As a result, a cheap rocket with a poor mass fraction may be able to deliver more payload to orbit with a given amount of money than a more complicated, more efficient rocket.

Launch assists

Many vehicles are only narrowly suborbital, so practically anything that gives a relatively small delta-v increase can be helpful, and outside assistance for a vehicle is therefore desirable.

Proposed launch assists include:

• sled launch (rail, maglev including Bantam, MagLifter, and StarTram, etc.)
• air launch or aircraft tow
• in-flight fueling
• Lofstrom launch loop/space fountains

And on-orbit resources such as:

• Space tether
• tugs

Nuclear propulsion

Main article: Nuclear propulsion

Due to weight issues such as shielding, many nuclear propulsion systems are unable to lift their own weight, and hence are unsuitable for launching to orbit. However, some designs such as the Orion project and some nuclear thermal designs do have a thrust to weight ratio in excess of 1, enabling them to lift off. Clearly, one of the main issues with nuclear propulsion would be safety, both during a launch for the passengers, but also in case of a failure during launch. As of December 2021, no current program is attempting nuclear propulsion from Earth's surface.

Beam-powered propulsion

Main article: Beam-powered propulsion

Because they can be more energetic than the potential energy that chemical fuel allows for, some laser or microwave powered rocket concepts have the potential to launch vehicles into orbit, single stage. In practice, this area is not possible with current technology.

Design challenges inherent in SSTO

The design space constraints of SSTO vehicles were described by rocket design engineer Robert Truax:

Using similar technologies (i.e., the same propellants and structural fraction), a two-stage-to-orbit vehicle will always have a better payload-to-weight ratio than a single stage designed for the same mission, in most cases, a very much better [payload-to-weight ratio]. Only when the structural factor approaches zero [very little vehicle structure weight] does the payload/weight ratio of a single-stage rocket approach that of a two-stage. A slight miscalculation and the single-stage rocket winds up with no payload. To get any at all, technology needs to be stretched to the limit. Squeezing out the last drop of specific impulse, and shaving off the last pound, costs money and/or reduces reliability.

The Tsiolkovsky rocket equation expresses the maximum change in velocity any single rocket stage can achieve:

$\mathrm{\Delta }v=I_{\text{sp}}\cdot g_0\ln (MR)$

where:

$\mathrm{\Delta }v$ (delta-v) is the maximum change of velocity of the vehicle,

$I_{\text{sp}}$ is the propellant specific impulse,

$g_0$ is the standard gravity,

$MR$ is the vehicle mass ratio,

$\ln$ refers to the natural logarithm function.

The mass ratio of a vehicle is defined as a ratio the initial vehicle mass when fully loaded with propellants $\left(m_i\right)$ to the final vehicle mass $\left(m_f\right)$ after the burn:

$MR=\frac{m_i}{m_f}=\frac{m_p+m_s+m_{\text{pl}}}{m_s+m_{\text{pl}}}$

where:

$m_i$ is the initial vehicle mass or the gross liftoff weight $\left(GLOW\right)$,

$m_f$ is the final vehicle mass after the burn,

$m_s$ is the structural mass of vehicle,

$m_p$ is the propellant mass,

$m_{\text{pl}}$ is the payload mass.

The propellant mass fraction ($\zeta$) of a vehicle can be expressed solely as a function of the mass ratio:

$\zeta =\frac{m_p}{m_i}=\frac{m_i-m_f}{m_i}=1-\frac{m_f}{m_i}=1-\frac{1}{MR}=\frac{MR-1}{MR}$

The structural coefficient ($\lambda$) is a critical parameter in SSTO vehicle design. Structural efficiency of a vehicle is maximized as the structural coefficient approaches zero. The structural coefficient is defined as:

Comparison of growth factor sensitivity for Single-Stage-to-Orbit (SSTO) and restricted stage Two-Stage-to-Orbit (TSTO) vehicles. Based on a LEO mission of Delta v = 9.1 km/s and payload mass = 4500 kg for range of propellant Isp.

$\lambda =\frac{m_s}{m_p+m_s}=\frac{m_s}{m_i-m_{\text{pl}}}=\frac{\frac{m_s}{m_i}}{1-\frac{m_{\text{pl}}}{m_i}}$

The overall structural mass fraction $\left(\frac{m_s}{m_i}\right)$ can be expressed in terms of the structural coefficient:

$\frac{m_s}{m_i}=\lambda \left(1-\frac{m_{\text{pl}}}{m_i}\right)$

An additional expression for the overall structural mass fraction can be found by noting that the payload mass fraction $\left(\frac{m_{\text{pl}}}{m_i}\right)$, propellant mass fraction and structural mass fraction sum to one:

$1=\frac{m_{\text{pl}}}{m_i}+\frac{m_p}{m_i}+\frac{m_s}{m_i}=\frac{m_{\text{pl}}}{m_i}+\zeta +\frac{m_s}{m_i}$

$\frac{m_s}{m_i}=1-\zeta -\frac{m_{\text{pl}}}{m_i}$

Equating the expressions for structural mass fraction and solving for the initial vehicle mass yields:

$m_i=GLOW=\frac{m_{\text{pl}}}{1-\left(\frac{\zeta }{1-\lambda }\right)}$

This expression shows how the size of a SSTO vehicle is dependent on its structural efficiency. Given a mission profile $\left(\mathrm{\Delta }v,m_{\text{pl}}\right)$ and propellant type $\left(I_{\text{sp}}\right)$, the size of a vehicle increases with an increasing structural coefficient. This growth factor sensitivity is shown parametrically for both SSTO and two-stage-to-orbit (TSTO) vehicles for a standard LEO mission. The curves vertically asymptote at the maximum structural coefficient limit where mission criteria can no longer be met:

${\lambda }_{\text{max}}=1-\zeta =\frac{1}{MR}$

In comparison to a non-optimized TSTO vehicle using restricted staging, a SSTO rocket launching an identical payload mass and using the same propellants will always require a substantially smaller structural coefficient to achieve the same delta-v. Given that current materials technology places a lower limit of approximately 0.1 on the smallest structural coefficients attainable, reusable SSTO vehicles are typically an impractical choice even when using the highest performance propellants available.

Examples

It is easier to achieve SSTO from a body with lower gravitational pull than Earth, such as the Moon or Mars. The Apollo Lunar Module ascended from the lunar surface to lunar orbit in a single stage.

A detailed study into SSTO vehicles was prepared by Chrysler Corporation's Space Division in 1970–1971 under NASA contract NAS8-26341. Their proposal (Shuttle SERV) was an enormous vehicle with more than 50,000 kilograms (110,000 lb) of payload, utilizing jet engines for (vertical) landing. While the technical problems seemed to be solvable, the USAF required a winged design that led to the Shuttle as we know it today.

The uncrewed DC-X technology demonstrator, originally developed by McDonnell Douglas for the Strategic Defense Initiative (SDI) program office, was an attempt to build a vehicle that could lead to an SSTO vehicle. The one-third-size test craft was operated and maintained by a small team of three people based out of a trailer, and the craft was once relaunched less than 24 hours after landing. Although the test program was not without mishap (including a minor explosion), the DC-X demonstrated that the maintenance aspects of the concept were sound. That project was cancelled when it landed with three of four legs deployed, tipped over, and exploded on the fourth flight after transferring management from the Strategic Defense Initiative Organization to NASA.

The Aquarius Launch Vehicle was designed to bring bulk materials to orbit as cheaply as possible.

Current development

Current and previous SSTO projects include the Japanese Kankoh-maru project, ARCA Haas 2C, Radian One and the Indian Avatar spaceplane.

Skylon

Main article: Reaction Engines Skylon

The British Government partnered with the ESA in 2010 to promote a single-stage to orbit spaceplane concept called Skylon. This design was pioneered by Reaction Engines Limited (REL), a company founded by Alan Bond after HOTOL was canceled. The Skylon spaceplane has been positively received by the British government, and the British Interplanetary Society. Following a successful propulsion system test that was audited by ESA's propulsion division in mid-2012, REL announced that it would begin a three-and-a-half-year project to develop and build a test jig of the Sabre engine to prove the engines performance across its air-breathing and rocket modes. In November 2012, it was announced that a key test of the engine precooler had been successfully completed, and that ESA had verified the precooler's design. The project's development is now allowed to advance to its next phase, which involves the construction and testing of a full-scale prototype engine.

Starship

Main article: SpaceX Starship

Elon Musk, CEO of SpaceX, has claimed that the upper stage of the prototype "Starship" rocket, currently in development in Starbase (Texas), has the capability to reach orbit as an SSTO. However he concedes that if this was done, there would be no appreciable mass left for a heat shield, landing legs, or fuel to land, much less any usable payload.

Alternative approaches to inexpensive spaceflight

Many studies have shown that regardless of selected technology, the most effective cost reduction technique is economies of scale. Merely launching a large total number reduces the manufacturing costs per vehicle, similar to how the mass production of automobiles brought about great increases in affordability.

Using this concept, some aerospace analysts believe the way to lower launch costs is the exact opposite of SSTO. Whereas reusable SSTOs would reduce per launch costs by making a reusable high-tech vehicle that launches frequently with low maintenance, the "mass production" approach views the technical advances as a source of the cost problem in the first place. By simply building and launching large quantities of rockets, and hence launching a large volume of payload, costs can be brought down. This approach was attempted in the late 1970s, early 1980s in West Germany with the Democratic Republic of the Congo-based OTRAG rocket.

This is somewhat similar to the approach some previous systems have taken, using simple engine systems with "low-tech" fuels, as the Russian and Chinese space programs still do.

An alternative to scale is to make the discarded stages practically reusable: this was the original design goal of the Space Shuttle phase B studies, and is currently pursued by the SpaceX reusable launch system development program with their Falcon 9, Falcon Heavy, and Starship, and Blue Origin using New Glenn.

Scramjet

From Wikipedia, the free encyclopedia

https://en.wikipedia.org/wiki/Scramjet 

A scramjet (supersonic combustion ramjet) is a variant of a ramjet airbreathing jet engine in which combustion takes place in supersonic airflow. As in ramjets, a scramjet relies on high vehicle speed to compress the incoming air forcefully before combustion (hence ramjet), but whereas a ramjet decelerates the air to subsonic velocities before combustion using shock cones, a scramjet has no shock cone and slows the airflow using shockwaves produced by its ignition source in place of a shock cone. This allows the scramjet to operate efficiently at extremely high speeds.

History

Before 2000

The Bell X-1 attained supersonic flight in 1947 and, by the early 1960s, rapid progress toward faster aircraft suggested that operational aircraft would be flying at "hypersonic" speeds within a few years. Except for specialized rocket research vehicles like the North American X-15 and other rocket-powered spacecraft, aircraft top speeds have remained level, generally in the range of Mach 1 to Mach 3.

During the US aerospaceplane program, between the 1950s and 1960s, Alexander Kartveli and Antonio Ferri were proponents of the scramjet approach.

In the 1950s and 1960s, a variety of experimental scramjet engines were built and ground tested in the US and the UK. Antonio Ferri successfully demonstrated a scramjet producing net thrust in November 1964, eventually producing 517 pounds-force (2.30 kN), about 80% of his goal. In 1958, an analytical paper discussed the merits and disadvantages of supersonic combustion ramjets. In 1964, Frederick S. Billig and Gordon L. Dugger submitted a patent application for a supersonic combustion ramjet based on Billig's PhD thesis. This patent was issued in 1981 following the removal of an order of secrecy.

In 1981, tests were made in Australia under the guidance of Professor Ray Stalker in the T3 ground test facility at ANU.

The first successful flight test of a scramjet was performed as a joint effort with NASA, over the Soviet Union in 1991. It was an axisymmetric hydrogen-fueled dual-mode scramjet developed by Central Institute of Aviation Motors (CIAM), Moscow in the late 1970s, but modernized with a FeCrAl alloy on a converted SM-6 missile to achieve initial flight parameters of Mach 6.8, before the scramjet flew at Mach 5.5. The scramjet flight was flown captive-carry atop the SA-5 surface-to-air missile that included an experimental flight support unit known as the "Hypersonic Flying Laboratory" (HFL), "Kholod".

Then, from 1992 to 1998, an additional six flight tests of the axisymmetric high-speed scramjet-demonstrator were conducted by CIAM together with France and then with NASA. Maximum flight velocity greater than Mach 6.4 was achieved and scramjet operation during 77 seconds was demonstrated. These flight test series also provided insight into autonomous hypersonic flight controls.

Progress in the 2000s

Main article: Scramjet programs

Artist's conception of the NASA X-43 with scramjet attached to the underside

In the 2000s, significant progress was made in the development of hypersonic technology, particularly in the field of scramjet engines.

The HyShot project demonstrated scramjet combustion on 30 July 2002. The scramjet engine worked effectively and demonstrated supersonic combustion in action. However, the engine was not designed to provide thrust to propel a craft. It was designed more or less as a technology demonstrator.

A joint British and Australian team from UK defense company Qinetiq and the University of Queensland were the first group to demonstrate a scramjet working in an atmospheric test.

Hyper-X claimed the first flight of a thrust-producing scramjet-powered vehicle with full aerodynamic maneuvering surfaces in 2004 with the X-43A. The last of the three X-43A scramjet tests achieved Mach 9.6 for a brief time.

On 15 June 2007, the US Defense Advanced Research Project Agency (DARPA), in cooperation with the Australian Defence Science and Technology Organisation (DSTO), announced a successful scramjet flight at Mach 10 using rocket engines to boost the test vehicle to hypersonic speeds.

A series of scramjet ground tests was completed at NASA Langley Arc-Heated Scramjet Test Facility (AHSTF) at simulated Mach 8 flight conditions. These experiments were used to support HIFiRE flight 2.

On 22 May 2009, Woomera hosted the first successful test flight of a hypersonic aircraft in HIFiRE (Hypersonic International Flight Research Experimentation). The launch was one of ten planned test flights. The series of flights is part of a joint research program between the Defence Science and Technology Organisation and the US Air Force, designated as the HIFiRE. HIFiRE is investigating hypersonics technology and its application to advanced scramjet-powered space launch vehicles; the objective is to support the new Boeing X-51 scramjet demonstrator while also building a strong base of flight test data for quick-reaction space launch development and hypersonic "quick-strike" weapons.

Progress in the 2010s

On 22 and 23 March 2010, Australian and American defense scientists successfully tested a (HIFiRE) hypersonic rocket. It reached an atmospheric velocity of "more than 5,000 kilometres per hour" (Mach 4) after taking off from the Woomera Test Range in outback South Australia.

On 27 May 2010, NASA and the United States Air Force successfully flew the X-51A Waverider for approximately 200 seconds at Mach 5, setting a new world record for flight duration at hypersonic airspeed. The Waverider flew autonomously before losing acceleration for an unknown reason and destroying itself as planned. The test was declared a success. The X-51A was carried aboard a B-52, accelerated to Mach 4.5 via a solid rocket booster, and then ignited the Pratt & Whitney Rocketdyne scramjet engine to reach Mach 5 at 70,000 feet (21,000 m). However, a second flight on 13 June 2011 was ended prematurely when the engine lit briefly on ethylene but failed to transition to its primary JP-7 fuel, failing to reach full power.

On 16 November 2010, Australian scientists from the University of New South Wales at the Australian Defence Force Academy successfully demonstrated that the high-speed flow in a naturally non-burning scramjet engine can be ignited using a pulsed laser source.

A further X-51A Waverider test failed on 15 August 2012. The attempt to fly the scramjet for a prolonged period at Mach 6 was cut short when, only 15 seconds into the flight, the X-51A craft lost control and broke apart, falling into the Pacific Ocean north-west of Los Angeles. The cause of the failure was blamed on a faulty control fin.

In May 2013, an uncrewed X-51A Waverider reached 4828 km/h (Mach 3.9) during a three-minute flight under scramjet power. The WaveRider was dropped at 50,000 feet (15,000 m) from a B-52 bomber, and then accelerated to Mach 4.8 by a solid rocket booster which then separated before the WaveRider's scramjet engine came into effect.

On 28 August 2016, the Indian space agency ISRO conducted a successful test of a scramjet engine on a two-stage, solid-fueled rocket. Twin scramjet engines were mounted on the back of the second stage of a two-stage, solid-fueled sounding rocket called Advanced Technology Vehicle (ATV), which is ISRO's advanced sounding rocket. The twin scramjet engines were ignited during the second stage of the rocket when the ATV achieved a speed of 7350 km/h (Mach 6) at an altitude of 20 km. The scramjet engines were fired for a duration of about 5 seconds.

On 12 June 2019, India successfully conducted the maiden flight test of its indigenously developed uncrewed scramjet demonstration aircraft for hypersonic speed flight from a base from Abdul Kalam Island in the Bay of Bengal at about 11:25 am. The aircraft is called the Hypersonic Technology Demonstrator Vehicle. The trial was carried out by the Defence Research and Development Organisation. The aircraft forms an important component of the country's programme for development of a hypersonic cruise missile system.

Progress in the 2020s

On 27 September 2021, DARPA announced successful flight of its Hypersonic Air-breathing Weapon Concept scramjet cruise missile. Another successful test was carried out in mid-March 2022 amid the Russian invasion of Ukraine. Details were kept secret to avoid escalating tension with Russia, only to be revealed by an unnamed Pentagon official in early April.

Design principles

Scramjet engines are a type of jet engine, and rely on the combustion of fuel and an oxidizer to produce thrust. Similar to conventional jet engines, scramjet-powered aircraft carry the fuel on board, and obtain the oxidizer by the ingestion of atmospheric oxygen (as compared to rockets, which carry both fuel and an oxidizing agent). This requirement limits scramjets to suborbital atmospheric propulsion, where the oxygen content of the air is sufficient to maintain combustion.

The scramjet is composed of three basic components: a converging inlet, where incoming air is compressed; a combustor, where gaseous fuel is burned with atmospheric oxygen to produce heat; and a diverging nozzle, where the heated air is accelerated to produce thrust. Unlike a typical jet engine, such as a turbojet or turbofan engine, a scramjet does not use rotating, fan-like components to compress the air; rather, the achievable speed of the aircraft moving through the atmosphere causes the air to compress within the inlet. As such, no moving parts are needed in a scramjet. In comparison, typical turbojet engines require multiple stages of rotating compressor rotors, and multiple rotating turbine stages, all of which add weight, complexity, and a greater number of failure points to the engine.

Due to the nature of their design, scramjet operation is limited to near-hypersonic velocities. As they lack mechanical compressors, scramjets require the high kinetic energy of a hypersonic flow to compress the incoming air to operational conditions. Thus, a scramjet-powered vehicle must be accelerated to the required velocity (usually about Mach 4) by some other means of propulsion, such as turbojet, railgun, or rocket engines. In the flight of the experimental scramjet-powered Boeing X-51A, the test craft was lifted to flight altitude by a Boeing B-52 Stratofortress before being released and accelerated by a detachable rocket to near Mach 4.5. In May 2013, another flight achieved an increased speed of Mach 5.1.

While scramjets are conceptually simple, actual implementation is limited by extreme technical challenges. Hypersonic flight within the atmosphere generates immense drag, and temperatures found on the aircraft and within the engine can be much greater than that of the surrounding air. Maintaining combustion in the supersonic flow presents additional challenges, as the fuel must be injected, mixed, ignited, and burned within milliseconds. While scramjet technology has been under development since the 1950s, only very recently have scramjets successfully achieved powered flight.

The compression, combustion, and expansion regions of: (a) turbojet, (b) ramjet, and (c) scramjet engines.

Basic principles

Scramjets are designed to operate in the hypersonic flight regime, beyond the reach of turbojet engines, and, along with ramjets, fill the gap between the high efficiency of turbojets and the high speed of rocket engines. Turbomachinery-based engines, while highly efficient at subsonic speeds, become increasingly inefficient at transonic speeds, as the compressor rotors found in turbojet engines require subsonic speeds to operate. While the flow from transonic to low supersonic speeds can be decelerated to these conditions, doing so at supersonic speeds results in a tremendous increase in temperature and a loss in the total pressure of the flow. Around Mach 3–4, turbomachinery is no longer useful, and ram-style compression becomes the preferred method.

Ramjets use high-speed characteristics of air to literally 'ram' air through an inlet diffuser into the combustor. At transonic and supersonic flight speeds, the air upstream of the inlet is not able to move out of the way quickly enough, and is compressed within the diffuser before being diffused into the combustor. Combustion in a ramjet takes place at subsonic velocities, similar to turbojets but the combustion products are then accelerated through a convergent-divergent nozzle to supersonic speeds. As they have no mechanical means of compression, ramjets cannot start from a standstill, and generally do not achieve sufficient compression until supersonic flight. The lack of intricate turbomachinery allows ramjets to deal with the temperature rise associated with decelerating a supersonic flow to subsonic speeds. However as speed rises, the internal energy of the flow after diffusor grows rapidly, so the relative addition of energy due to fuel combustion becomes lower, leading to decrease in efficiency of the engine. This leads to decrease in thrust generated by ramjets at higher speeds. 

Thus, to generate thrust at very high velocities, rise of pressure and temperature of incoming air flow must be tightly controlled. In particular, this means that deceleration of the airflow to subsonic speed cannot be allowed. At such setup not only mixing of fuel with air presents an engineering challenge, but also speed of combustion in air-fuel mixtures becomes a concern. In addition the relative increase of internal energy in combustion chamber must be maximized. Consequently, current scramjet technology requires the use of high-energy fuels and active cooling schemes to maintain sustained operation, often using hydrogen and regenerative cooling techniques.

Theory

All scramjet engines have an intake which compresses the incoming air, fuel injectors, a combustion chamber, and a divergent thrust nozzle. Sometimes engines also include a region which acts as a flame holder, although the high stagnation temperatures mean that an area of focused waves may be used, rather than a discrete engine part as seen in turbine engines. Other engines use pyrophoric fuel additives, such as silane, to avoid flameout. An isolator between the inlet and combustion chamber is often included to improve the homogeneity of the flow in the combustor and to extend the operating range of the engine.

Shockwave imaging by the University of Maryland using Schlieren imaging determined that the fuel mixture controls compression by creating backpressure and shockwaves that slow and compress the air before ignition, much like the shock cone of a Ramjet. The imaging showed that the higher the fuel flow and combustion, the more shockwaves formed ahead of the combustor, which slowed and compressed the air before ignition.

Computational fluid dynamics (CFD) image of the NASA X-43A with scramjet attached to the underside at Mach 7

A scramjet is reminiscent of a ramjet. In a typical ramjet, the supersonic inflow of the engine is decelerated at the inlet to subsonic speeds and then reaccelerated through a nozzle to supersonic speeds to produce thrust. This deceleration, which is produced by a normal shock, creates a total pressure loss which limits the upper operating point of a ramjet engine.

For a scramjet, the kinetic energy of the freestream air entering the scramjet engine is largely comparable to the energy released by the reaction of the oxygen content of the air with a fuel (e.g. hydrogen). Thus the heat released from combustion at Mach 2.5 is around 10% of the total enthalpy of the working fluid. Depending on the fuel, the kinetic energy of the air and the potential combustion heat release will be equal at around Mach 8. Thus the design of a scramjet engine is as much about minimizing drag as maximizing thrust.

This high speed makes the control of the flow within the combustion chamber more difficult. Since the flow is supersonic, no downstream influence propagates within the freestream of the combustion chamber. Throttling of the entrance to the thrust nozzle is not a usable control technique. In effect, a block of gas entering the combustion chamber must mix with fuel and have sufficient time for initiation and reaction, all the while traveling supersonically through the combustion chamber, before the burned gas is expanded through the thrust nozzle. This places stringent requirements on the pressure and temperature of the flow, and requires that the fuel injection and mixing be extremely efficient. Usable dynamic pressures lie in the range 20 to 200 kilopascals (2.9 to 29.0 psi), where

$q=\frac{1}{2}\rho v^2$

where

q is the dynamic pressure of the gas

ρ (rho) is the density of the gas

v is the velocity of the gas

To keep the combustion rate of the fuel constant, the pressure and temperature in the engine must also be constant. This is problematic because the airflow control systems that would facilitate this are not physically possible in a scramjet launch vehicle due to the large speed and altitude range involved, meaning that it must travel at an altitude specific to its speed. Because air density reduces at higher altitudes, a scramjet must climb at a specific rate as it accelerates to maintain a constant air pressure at the intake. This optimal climb/descent profile is called a "constant dynamic pressure path". It is thought that scramjets might be operable up to an altitude of 75 km.

Fuel injection and management is also potentially complex. One possibility would be that the fuel be pressurized to 100 bar by a turbo pump, heated by the fuselage, sent through the turbine and accelerated to higher speeds than the air by a nozzle. The air and fuel stream are crossed in a comb-like structure, which generates a large interface. Turbulence due to the higher speed of the fuel leads to additional mixing. Complex fuels like kerosene need a long engine to complete combustion.

The minimum Mach number at which a scramjet can operate is limited by the fact that the compressed flow must be hot enough to burn the fuel, and have pressure high enough that the reaction be finished before the air moves out the back of the engine. Additionally, to be called a scramjet, the compressed flow must still be supersonic after combustion. Here two limits must be observed: First, since when a supersonic flow is compressed it slows down, the level of compression must be low enough (or the initial speed high enough) not to slow the gas below Mach 1. If the gas within a scramjet goes below Mach 1 the engine will "choke", transitioning to subsonic flow in the combustion chamber. This effect is well known amongst experimenters on scramjets since the waves caused by choking are easily observable. Additionally, the sudden increase in pressure and temperature in the engine can lead to an acceleration of the combustion, leading to the combustion chamber exploding.

Second, the heating of the gas by combustion causes the speed of sound in the gas to increase (and the Mach number to decrease) even though the gas is still travelling at the same speed. Forcing the speed of air flow in the combustion chamber under Mach 1 in this way is called "thermal choking". It is clear that a pure scramjet can operate at Mach numbers of 6–8, but in the lower limit, it depends on the definition of a scramjet. There are engine designs where a ramjet transforms into a scramjet over the Mach 3–6 range, known as dual-mode scramjets. In this range however, the engine is still receiving significant thrust from subsonic combustion of the ramjet type.

The high cost of flight testing and the unavailability of ground facilities have hindered scramjet development. A large amount of the experimental work on scramjets has been undertaken in cryogenic facilities, direct-connect tests, or burners, each of which simulates one aspect of the engine operation. Further, vitiated facilities (with the ability to control air impurities), storage heated facilities, arc facilities and the various types of shock tunnels each have limitations which have prevented perfect simulation of scramjet operation. The HyShot flight test showed the relevance of the 1:1 simulation of conditions in the T4 and HEG shock tunnels, despite having cold models and a short test time. The NASA-CIAM tests provided similar verification for CIAM's C-16 V/K facility and the Hyper-X project is expected to provide similar verification for the Langley AHSTF, CHSTF, and 8 ft (2.4 m) HTT.

Computational fluid dynamics has only recently reached a position to make reasonable computations in solving scramjet operation problems. Boundary layer modeling, turbulent mixing, two-phase flow, flow separation, and real-gas aerothermodynamics continue to be problems on the cutting edge of CFD. Additionally, the modeling of kinetic-limited combustion with very fast-reacting species such as hydrogen makes severe demands on computing resources. Reaction schemes are numerically stiff requiring reduced reaction schemes.

Much of scramjet experimentation remains classified. Several groups, including the US Navy with the SCRAM engine between 1968 and 1974, and the Hyper-X program with the X-43A, have claimed successful demonstrations of scramjet technology. Since these results have not been published openly, they remain unverified and a final design method of scramjet engines still does not exist.

The final application of a scramjet engine is likely to be in conjunction with engines which can operate outside the scramjet's operating range. Dual-mode scramjets combine subsonic combustion with supersonic combustion for operation at lower speeds, and rocket-based combined cycle (RBCC) engines supplement a traditional rocket's propulsion with a scramjet, allowing for additional oxidizer to be added to the scramjet flow. RBCCs offer a possibility to extend a scramjet's operating range to higher speeds or lower intake dynamic pressures than would otherwise be possible.

Advantages and disadvantages of scramjets

Advantages

1. Does not have to carry oxygen
2. No rotating parts makes it easier to manufacture than a turbojet
3. Has a higher specific impulse (change in momentum per unit of propellant) than a rocket engine; could provide between 1000 and 4000 seconds, while a rocket typically provides around 450 seconds or less.
4. Higher speed could mean cheaper access to outer space in the future

Disadvantages

1. Difficult / expensive testing and development
2. Very high initial propulsion requirements

Special cooling and materials

Unlike a rocket that quickly passes mostly vertically through the atmosphere or a turbojet or ramjet that flies at much lower speeds, a hypersonic airbreathing vehicle optimally flies a "depressed trajectory", staying within the atmosphere at hypersonic speeds. Because scramjets have only mediocre thrust-to-weight ratios, acceleration would be limited. Therefore, time in the atmosphere at supersonic speed would be considerable, possibly 15–30 minutes. Similar to a reentering space vehicle, heat insulation would be a formidable task, with protection required for a duration longer than that of a typical space capsule, although less than the Space Shuttle.

New materials offer good insulation at high temperature, but they often sacrifice themselves in the process. Therefore, studies often plan on "active cooling", where coolant circulating throughout the vehicle skin prevents it from disintegrating. Often the coolant is the fuel itself, in much the same way that modern rockets use their own fuel and oxidizer as coolant for their engines. All cooling systems add weight and complexity to a launch system. The cooling of scramjets in this way may result in greater efficiency, as heat is added to the fuel prior to entry into the engine, but results in increased complexity and weight which ultimately could outweigh any performance gains.

Vehicle performance

The specific impulse of various engines

The performance of a launch system is complex and depends greatly on its weight. Normally craft are designed to maximise range ($R$), orbital radius ($R$) or payload mass fraction ($\mathrm{\Gamma }$) for a given engine and fuel. This results in tradeoffs between the efficiency of the engine (takeoff fuel weight) and the complexity of the engine (takeoff dry weight), which can be expressed by the following:

${\displaystyle {\mathrm{\Pi }}_{\text{e}}+{\mathrm{\Pi }}_{\text{f}}+\frac{1}{\mathrm{\Gamma }}=1}$

Where :

• ${\displaystyle {\mathrm{\Pi }}_{\text{e}}=\frac{m_{\text{empty}}}{m_{\text{initial}}}}$ is the empty mass fraction, and represents the weight of the superstructure, tankage and engine.
• ${\displaystyle {\mathrm{\Pi }}_{\text{f}}=\frac{m_{\text{fuel}}}{m_{\text{initial}}}}$ is the fuel mass fraction, and represents the weight of fuel, oxidiser and any other materials which are consumed during the launch.
• ${\displaystyle \mathrm{\Gamma }=\frac{m_{\text{initial}}}{m_{\text{payload}}}}$ is initial mass ratio, and is the inverse of the payload mass fraction. This represents how much payload the vehicle can deliver to a destination.

A scramjet increases the mass of the motor ${\displaystyle {\mathrm{\Pi }}_{\text{e}}}$ over a rocket, and decreases the mass of the fuel ${\displaystyle {\mathrm{\Pi }}_{\text{f}}}$. It can be difficult to decide whether this will result in an increased $\mathrm{\Gamma }$ (which would be an increased payload delivered to a destination for a constant vehicle takeoff weight). The logic behind efforts driving a scramjet is (for example) that the reduction in fuel decreases the total mass by 30%, while the increased engine weight adds 10% to the vehicle total mass. Unfortunately the uncertainty in the calculation of any mass or efficiency changes in a vehicle is so great that slightly different assumptions for engine efficiency or mass can provide equally good arguments for or against scramjet powered vehicles.

Additionally, the drag of the new configuration must be considered. The drag of the total configuration can be considered as the sum of the vehicle drag ($D$) and the engine installation drag (${\displaystyle D_{\text{e}}}$). The installation drag traditionally results from the pylons and the coupled flow due to the engine jet, and is a function of the throttle setting. Thus it is often written as:

${\displaystyle D_{\text{e}}={\varphi }_{\text{e}}F}$

Where:

• ${\displaystyle {\varphi }_{\text{e}}}$ is the loss coefficient
• $F$ is the thrust of the engine

For an engine strongly integrated into the aerodynamic body, it may be more convenient to think of (${\displaystyle D_{\text{e}}}$) as the difference in drag from a known base configuration.

The overall engine efficiency can be represented as a value between 0 and 1 (${\eta }_0$), in terms of the specific impulse of the engine:

${\displaystyle {\eta }_0=\frac{g_0{V}_0}{h_{\text{PR}}}{I}_{\text{sp}}=\frac{\text{Thrust power}}{\text{Chemical energy rate}}}$

Where:

• $g_0$ is the acceleration due to gravity at ground level
• $V_0$ is the vehicle speed
• $I_{\text{sp}}$ is the specific impulse
• ${\displaystyle h_{\text{PR}}}$ is fuel heat of reaction

Specific impulse is often used as the unit of efficiency for rockets, since in the case of the rocket, there is a direct relation between specific impulse, specific fuel consumption and exhaust velocity. This direct relation is not generally present for airbreathing engines, and so specific impulse is less used in the literature. Note that for an airbreathing engine, both ${\eta }_0$ and $I_{\text{sp}}$ are a function of velocity.

The specific impulse of a rocket engine is independent of velocity, and common values are between 200 and 600 seconds (450 s for the space shuttle main engines). The specific impulse of a scramjet varies with velocity, reducing at higher speeds, starting at about 1200 s, although values in the literature vary.

For the simple case of a single stage vehicle, the fuel mass fraction can be expressed as:

${\displaystyle {\mathrm{\Pi }}_{\text{f}}=1-\exp \left[-\frac{\left(\frac{V_{\text{initial}}^2}{2}-\frac{V_i^2}{2}\right)+\int gdr}{{\eta }_0{h}_{\text{PR}}\left(1-\frac{D+D_{\text{e}}}{F}\right)}\right]}$

Where this can be expressed for single stage transfer to orbit as:

${\displaystyle {\mathrm{\Pi }}_{\text{f}}=1-\exp \left[-\frac{g_0{r}_0\left(1-\frac{1}{2}\frac{r_0}{r}\right)}{{\eta }_0{h}_{\text{PR}}\left(1-\frac{D+D_{\text{e}}}{F}\right)}\right]}$

or for level atmospheric flight from air launch (missile flight):

${\displaystyle {\mathrm{\Pi }}_{\text{f}}=1-\exp \left[-\frac{g_{0}R}{{\eta }_0{h}_{\text{PR}}\left(1-{\varphi }_{\text{e}}\right)\frac{C_{\text{L}}}{C_{\text{D}}}}\right]}$

Where $R$ is the range, and the calculation can be expressed in the form of the Breguet range formula:

${\displaystyle \begin{array}{rl}{\mathrm{\Pi }}_{\text{f}}& =1-e^{-BR}\\ B& =\frac{g_0}{{\eta }_0{h}_{PR}\left(1-{\varphi }_e\right)\frac{C_{\text{L}}}{C_{\text{D}}}}\end{array}}$

Where:

• ${\displaystyle C_{\text{L}}}$ is the lift coefficient
• ${\displaystyle C_{\text{D}}}$ is the drag coefficient

This extremely simple formulation, used for the purposes of discussion assumes:

• Single stage vehicle
• No aerodynamic lift for the transatmospheric lifter

However they are true generally for all engines.

Initial propulsion requirements

A scramjet cannot produce efficient thrust unless boosted to high speed, around Mach 5, although depending on the design it could act as a ramjet at low speeds. A horizontal take-off aircraft would need conventional turbofan, turbojet, or rocket engines to take off, sufficiently large to move a heavy craft. Also needed would be fuel for those engines, plus all engine-associated mounting structure and control systems. Turbofan and turbojet engines are heavy and cannot easily exceed about Mach 2–3, so another propulsion method would be needed to reach scramjet operating speed. That could be ramjets or rockets. Those would also need their own separate fuel supply, structure, and systems. Many proposals instead call for a first stage of droppable solid rocket boosters, which greatly simplifies the design.

Testing difficulties

Test of Pratt & Whitney Rocketdyne SJY61 scramjet engine for the Boeing X-51

Unlike jet or rocket propulsion systems facilities which can be tested on the ground, testing scramjet designs uses extremely expensive hypersonic test chambers or expensive launch vehicles, both of which lead to high instrumentation costs. Tests using launched test vehicles very typically end with destruction of the test item and instrumentation.

Advantages and disadvantages for orbital vehicles

Propellant

An advantage of a hypersonic airbreathing (typically scramjet) vehicle like the X-30 is avoiding or at least reducing the need for carrying oxidizer. For example, the Space Shuttle external tank held 616,432.2 kg of liquid oxygen (LOX) and 103,000 kg of liquid hydrogen (LH₂) while having an empty weight of 30,000 kg. The orbiter gross weight was 109,000 kg with a maximum payload of about 25,000 kg and to get the assembly off the launch pad the shuttle used two very powerful solid rocket boosters with a weight of 590,000 kg each. If the oxygen could be eliminated, the vehicle could be lighter at liftoff and possibly carry more payload.

On the other hand, scramjets spend more time in the atmosphere and require more hydrogen fuel to deal with aerodynamic drag. Whereas liquid oxygen is quite a dense fluid (1141 kg/m³), liquid hydrogen has much lower density (70.85 kg/m³) and takes up more volume. This means that the vehicle using this fuel becomes much bigger and gives more drag. Other fuels have more comparable density, such as RP-1 (810 kg/m³) JP-7 (density at 15 °C 779–806 kg/m³) and unsymmetrical dimethylhydrazine (UDMH) (793.00 kg/m³).

Thrust-to-weight ratio

One issue is that scramjet engines are predicted to have exceptionally poor thrust-to-weight ratio of around 2, when installed in a launch vehicle. A rocket has the advantage that its engines have very high thrust-weight ratios (~100:1), while the tank to hold the liquid oxygen approaches a volume ratio of ~100:1 also. Thus a rocket can achieve a very high mass fraction, which improves performance. By way of contrast the projected thrust/weight ratio of scramjet engines of about 2 mean a much larger percentage of the takeoff mass is engine (ignoring that this fraction increases anyway by a factor of about four due to the lack of onboard oxidiser). In addition the vehicle's lower thrust does not necessarily avoid the need for the expensive, bulky, and failure-prone high performance turbopumps found in conventional liquid-fuelled rocket engines, since most scramjet designs seem to be incapable of orbital speeds in airbreathing mode, and hence extra rocket engines are needed.

Need for additional propulsion to reach orbit

Scramjets might be able to accelerate from approximately Mach 5–7 to around somewhere between half of orbital speed and orbital speed (X-30 research suggested that Mach 17 might be the limit compared to an orbital speed of Mach 25, and other studies put the upper speed limit for a pure scramjet engine between Mach 10 and 25, depending on the assumptions made). Generally, another propulsion system (very typically, a rocket is proposed) is expected to be needed for the final acceleration into orbit. Since the delta-V is moderate and the payload fraction of scramjets high, lower performance rockets such as solids, hypergolics, or simple liquid fueled boosters might be acceptable.

Theoretical projections place the top speed of a scramjet between Mach 12 (14,000 km/h; 8,400 mph) and Mach 24 (25,000 km/h; 16,000 mph). For comparison, the orbital speed at 200 kilometres (120 mi) low Earth orbit is 7.79 kilometres per second (28,000 km/h; 17,400 mph).

Reentry

The scramjet's heat-resistant underside potentially doubles as its reentry system if a single-stage-to-orbit vehicle using non-ablative, non-active cooling is visualised. If an ablative shielding is used on the engine it will probably not be usable after ascent to orbit. If active cooling is used with the fuel as coolant, the loss of all fuel during the burn to orbit will also mean the loss of all cooling for the thermal protection system.

Costs

Reducing the amount of fuel and oxidizer does not necessarily improve costs as rocket propellants are comparatively very cheap. Indeed, the unit cost of the vehicle can be expected to end up far higher, since aerospace hardware cost is about two orders of magnitude higher than liquid oxygen, fuel and tankage, and scramjet hardware seems to be much heavier than rockets for any given payload. Still, if scramjets enable reusable vehicles, this could theoretically be a cost benefit. Whether equipment subject to the extreme conditions of a scramjet can be reused sufficiently many times is unclear; all flown scramjet tests only survive for short periods and have never been designed to survive a flight to date.

The eventual cost of such a vehicle is the subject of intense debate since even the best estimates disagree whether a scramjet vehicle would be advantageous. It is likely that a scramjet vehicle would need to lift more load than a rocket of equal takeoff weight to be equally as cost efficient (if the scramjet is a non-reusable vehicle).

Issues

Space launch vehicles may or may not benefit from having a scramjet stage. A scramjet stage of a launch vehicle theoretically provides a specific impulse of 1000 to 4000 s whereas a rocket provides less than 450 s while in the atmosphere. A scramjet's specific impulse decreases rapidly with speed, however, and the vehicle would suffer from a relatively low lift to drag ratio.

The installed thrust to weight ratio of scramjets compares very unfavorably with the 50–100 of a typical rocket engine. This is compensated for in scramjets partly because the weight of the vehicle would be carried by aerodynamic lift rather than pure rocket power (giving reduced 'gravity losses'), but scramjets would take much longer to get to orbit due to lower thrust which greatly offsets the advantage. The takeoff weight of a scramjet vehicle is significantly reduced over that of a rocket, due to the lack of onboard oxidiser, but increased by the structural requirements of the larger and heavier engines.

Whether this vehicle could be reusable or not is still a subject of debate and research.

Proposed applications

An aircraft using this type of jet engine could dramatically reduce the time it takes to travel from one place to another, potentially putting any place on Earth within a 90-minute flight. However, there are questions about whether such a vehicle could carry enough fuel to make useful length trips. In addition, some countries ban or penalize airliners and other civil aircraft that create sonic booms. (For example, in the United States, FAA regulations prohibit supersonic flights over land, by civil aircraft.)

Scramjet vehicle has been proposed for a single stage to tether vehicle, where a Mach 12 spinning orbital tether would pick up a payload from a vehicle at around 100 km and carry it to orbit.