The Aircraft • It does not have to carry oxygen • No rotating parts makes it easier to manufacture than a turbojet • Has a higher
specific impulse (change in momentum per unit of propellant) than a rocket engine; could provide about 1000 seconds at Mach 7, while a rocket typically provides around 450 seconds or less. • Higher speed could mean cheaper access to outer space in the future • Difficult / expensive testing and development • Very high initial propulsion requirements 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. 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 (\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: :\Pi_\text{e} + \Pi_\text{f} + \frac{1}{\Gamma} = 1 Where : • \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. • \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. • \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 \Pi_\text{e} over a rocket, and decreases the mass of the fuel \Pi_\text{f}. It can be difficult to decide whether this will result in an increased \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 (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: :D_\text{e} = \phi_\text{e} F Where: • \phi_\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 (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: :\eta_0 = \frac{g_0 V_0}{h_\text{PR}} I_\text{sp} = \frac{\mbox{Thrust power}}{\mbox{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 • 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 (450s for the space shuttle main engines). The specific impulse of a scramjet varies with velocity, reducing at higher speeds, starting at about 1200s, although values in the literature vary. For the simple case of a single stage vehicle, the fuel mass fraction can be expressed as: :\Pi_\text{f} = 1 - \exp\left[-\frac{\left(\frac{V_\text{initial}^2}{2} - \frac{V_i^2}{2}\right) + \int{g}\,dr}{\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: :\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): :\Pi_\text{f} = 1 - \exp\left[-\frac{g_0 R}{\eta_0 h_\text{PR}\left(1 - \phi_\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: :\begin{align} \Pi_\text{f} &= 1 - e^{-BR} \\ B &= \frac{g_0}{\eta_0 h_{PR}\left(1 - \phi_e\right)\frac{C_\text{L}}{C_\text{D}}} \end{align} Where: • C_\text{L} is the
lift coefficient • 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. A scramjet cannot produce efficient thrust unless boosted to high speed, around Mach5, 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 Mach2–3, 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. A number of proposals instead call for a first stage of droppable
solid rocket boosters, which greatly simplifies the design.
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.
Orbital vehicles 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 (LH) 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). 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. Scramjets might be able to accelerate from approximately Mach5–7 to around somewhere between half of
orbital speed and orbital speed (X-30 research suggested that Mach17 might be the limit compared to an orbital speed of Mach25, and other studies put the upper speed limit for a pure scramjet engine between Mach10 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 and . For comparison, the orbital speed at
low Earth orbit is . 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. 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 multiple 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). 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 4000s whereas a rocket provides less than 450s 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==