Posted at 12.29.2018
This paper talks about the relevant selection standards for a single stage to orbit (SSTO) propulsion system and then reviews the characteristics of the typical engine types proposed because of this role against these criteria. The engine motor types considered include Hydrogen/Oxygen (H2/O2) rockets, Scramjets, Turbojets, Turborockets and Water Air Cycle Engines. In the writers opinion none of them of the above motors have the ability to meet all the necessary standards for an SSTO propulsion system concurrently. However by selecting appropriate features from each it is possible to synthesise a new class of motors that happen to be specifically optimised for the SSTO role. The resulting machines utilize precooling of the airstream and a high internal pressure ratio to enable a relatively conventional high pressure rocket combustion chamber to be utilised in both airbreathing and rocket methods. This brings about a substantial mass conserving with set up advantages which by careful design of the cycle thermodynamics enables the entire probable of airbreathing to be realised. The SABRE engine unit which forces the SKYLON introduction vehicle can be an example of one of the so called 'Precooled cross types airbreathing rocket motors' and the conceptual reasoning which causes its main design guidelines are described in the newspaper.
Keywords: Reusable launchers, SABRE, SKYLON, SSTO
Several organisations world-wide are studying the technical and commercial feasibility of reusable SSTO launchers. This new category of vehicles appear to provide tantalising potential client of greatly reduced continuing costs and increased stability compared to existing expendable vehicles. However achieving this breakthrough is a difficult task since the attainment of orbital speed in a re-entry in a position single stage calls for incredible propulsive performance.
Most studies at this point have focused on high pressure hydrogen/air (H2/O2) rocket engines for the primary propulsion of such vehicles. Nonetheless it is the writers thoughts and opinions that despite recent advancements in materials technology this procedure is not destined to achieve success, because of the relatively low specific impulse of the kind of propulsion. Airbreathing motors give a possible route frontward with their intrinsically higher specific impulse. However their low thrust/weight ratio, limited Mach number range and high vibrant pressure trajectory have before cancelled any theoretical benefits.
By design review of the relevant characteristics of both rockets and airbreathing motors this paper models out the rationale for selecting deeply precooled hybrid airbreathing rocket machines for the main propulsion system of SSTO launchers as exemplified by the SKYLON vehicle .
This paper will only consider those engine types which would result in politically and environmentally suitable vehicles. Therefore engines using nuclear reactions (eg: onboard fission reactors or external nuclear pulse) and chemical engines with dangerous exhausts (eg: fluorine/air) will be excluded.
The candidate motors can be split into two broad categories, namely 100 % pure rockets and motors with an airbreathing part. Since nothing of the airbreathers are capable of accelerating an SSTO vehicle completely to orbital velocity, a sensible vehicle will always have an onboard rocket engine unit to complete the ascent. Which means use of airbreathing has always been suggested within the context of improving the specific impulse of genuine rocket propulsion during the primary lower Mach part of the trajectory.
Airbreathing machines have a much lower thrust/ weight ratio than rocket engines (Л†10%) which will offset the good thing about reduced fuel ingestion. Therefore vehicles with airbreathing engines invariably have wings and employ a lifting trajectory to be able to lessen the installed thrust necessity and therefore the airbreathing engine unit mass charges. The blend of wings and airbreathing engines then demands a low smooth trajectory (compared to a ballistic rocket trajectory) to be able to increase the installed performance (i. e. (thrust-drag)/fuel movement). This high powerful pressure trajectory gives rise to one of the downsides associated with an airbreathing approach because the airframe heating system and loading are increased during the ascent which in the end displays in increased framework mass. Nevertheless the absolute level of mass growth depends upon the relative seriousness of the ascent in comparison with reentry which is mostly dependant on the type of airbreathing engine chosen. An additional drawback to the reduced trajectory is increased pull losses particularly because the vehicle loiters longer in the lower atmosphere due to the lower acceleration, offset somewhat by the much reduced gravity reduction through the rocket powered ascent.
Importantly however, the addition of a couple of wings brings more than just performance advantages to airbreathing vehicles. They also give significantly increased abort capability since an adequately configured vehicle can stay in stable airline flight with up to 1 / 2 of its propulsion systems shutdown. Also during reentry the presence of wings reduces the ballistic coefficient therefore reducing the home heating and hence thermal safety system mass, whilst concurrently improving the vehicle lift/drag proportion permitting greater crossrange.
The suitability of the following motors to the SSTO launcher role will be discussed since these are representative of the primary types currently under research within various organisations world-wide:
The selection of an 'optimum' propulsion system requires an diagnosis of a number of interdependant factors which can be the following. The relative importance of these factors depends on the severe nature of the quest and the automobile characteristics.
Useable Mach amount and altitude range.
Installed specific impulse.
Performance level of sensitivity to component level efficiencies.
Effect on airframe design (Cg/Cp pitch trim & structural efficiency).
Effect of required engine unit trajectory (Q and warming) on airframe technology/materials.
Materials/structures/aerothermodynamic and making technology.
Engine scale and technology level.
Complexity and ability demand of earth test facilities.
Necessity of your X plane research study to precede the key development program.
Hydrogen/air rocket motors achieve an extremely high thrust/weight ratio (60-80) but relatively low specific impulse (450-475 secs in vacuum) compared with conventional airbreathing engines. Due to the relatively large Л† V had a need to reach low globe orbit (approx 9 km/s including gravity and move losses) in relation to the engine unit exhaust velocity, SSTO rocket vehicles are characterised by very high mass ratios and low payload fractions.
The H2/O2 propellant mixture is invariably chosen for SSTO rockets due to its higher performance than other alternatives regardless of the structural fines of having a very low thickness cryogenic fuel. In order to maximise the precise impulse, high area ratio nozzles will be required which inevitably brings about a higher chamber pressure cycle in order to provide a compact assembly and reduce again pressure losses at low altitude. The need to minimise back pressure losses normally ends in selecting some type of altitude compensating nozzle since standard bell nozzles have high divergence and overexpansion losses when running in a segregated condition.
The high thrust/weight and low specific impulse of H2/O2 rocket motors favours vertical takeoff wingless vehicles because the wing mass and pull penalty of your lifting trajectory results a smaller payload than a steep ballistic climb out of the atmosphere. The ascent trajectory is therefore extremely harmless (in terms of vibrant pressure and warming) with vehicle materials selection dependant on re-entry. Relative to airbreathing vehicles a pure rocket vehicle has an increased density (gross remove weight/volume) because of the reduced hydrogen intake that includes a favourable effect on the tankage and thermal safety system mass.
In their favour rocket motors stand for broadly known (current) technology, are surface testable in simple facilities, efficient throughout the complete Mach amount range and bodily very compact resulting in good engine motor/airframe integration. Abort capabilities for an SSTO rocket vehicle would be performed by getting a high takeoff thrust/weight ratio (eg: 1. 5) and a big number of machines (eg: 10) to permit shutdown of at least two whilst retaining overall vehicle control. From an functional standpoint SSTO rockets will be relatively noisy because the high takeoff mass and thrust/weight proportion results in an installed thrust level up to 10 times higher than a well designed airbreather.
Reentry should be relatively straightforward providing the automobile reenters bottom part first with productive cooling down of the engine unit nozzles and the automobile base. Nevertheless the maximum lift up/drag ratio in this attitude is relatively low (approx 0. 25) limiting the maximum possible crossrange to around 250 kilometres. Having reached a low altitude a few of the main machines would be restarted to control the subsonic descent before finally effecting a tailfirst getting on thighs. Low crossrange is not really a particular problem providing the vehicle operator has enough time to wait for the orbital aircraft to cross the getting site. However in the case of your military services or commercial operator this may pose a significant operational restriction and is consequently regarded as an undesirable characteristic for a fresh unveiling vehicle.
In an effort to boost the crossrange ability some designs strive nosefirst re-entry of a blunt cone molded vehicle or additionally a combined wing/body configuration. This approach potentially escalates the lift/drag ratio by lowering the fuselage influx pull and/or increasing the aerodynamic lift up generation. However the drawback to the approach would be that the nosefirst attitude is aerodynamically unpredictable since the aft mounted engine motor package pulls the vacant center of gravity a significant distance behind the hypersonic middle of pressure. The causing pitching minute is difficult to lean without adding nose ballast or large control floors projecting from the automobile base. It is expected that the additional mass of the components will probably erode the small payload capacity for this engine motor/vehicle combination to the main point where it is no more feasible.
Recent advancements in materials technology (eg: fibre strengthened plastics and ceramics) have made a large impact on the feasibility of the vehicles. Nevertheless the payload fraction is still really small at around 1-2% for an Equatorial low Globe orbit falling to only 0. 25% for a Polar orbit. The low payload fraction is normally perceived to be the key disadvantage of this engine/vehicle combination and has historically prevented the introduction of such vehicles, since it is felt that a small degree of optimism in the initial mass estimates may be concealing the actual fact that the 'real' payload portion is negative.
One possible course onward to increasing the common specific impulse of rocket vehicles is to hire the atmosphere for both oxidiser and response mass for part of the ascent. That is an old idea dating back to the 1950's and revitalised by the emergence of the BAe/Rolls Royce 'HOTOL' task in the 1980's . The following portions will review the primary airbreathing engine applicants and trace the design qualifications of precooled hybrid airbreathing rockets.
A ramjet engine is from a thermodynamic point of view a very simple device consisting of an intake, combustion and nozzle system where the cycle pressure go up is achieved strictly by ram compression. Consequently a separate propulsion system is needed to accelerate the vehicle to speeds at which the ramjet can takeover (Mach 1-2). A conventional hydrogen fuelled ramjet with a subsonic combustor is capable of operating up to around Mach 5-6 at which point the restricting effects of dissociation decrease the effective temperature addition to the airflow producing a rapid reduction in nett thrust. The idea behind the scramjet engine unit is to enough time dissociation limit by only partly slowing the airstream through the intake system (thus reducing the static temps rise) and hence permitting greater useful temperature addition in the now supersonic combustor. By this implies scramjet engines offer the tantalising possibility of achieving a high specific impulse up to high Mach volumes. The consequent reduction in the rocket driven Л† V would translate into a large keeping in the mass of water oxygen required and hence possibly a reduction in start mass.
Although the scramjet is theoretically with the capacity of generating positive nett thrust to a significant portion of orbital velocity it is unworkable at low supersonic speeds. Therefore it is generally suggested that the internal geometry be reconfigured to function as a conventional ramjet to Mach 5 followed by change to scramjet function. A further reduced amount of the useful velocity selection of the scramjet results from awareness of the nett vehicle specific impulse ((thrust-drag)/fuel circulation) in scramjet function as compared with rocket mode. This tradeoff implies that it works more effectively to shut the scramjet down at Mach 12-15 and continue the remainder of the ascent on natural rocket vitality. Therefore a scramjet powered launcher could have four main propulsion methods: a minimal speed accelerator method to ramjet followed by scramjet and lastly rocket method. The proposed low velocity propulsor is usually a ducted ejector rocket system using the scramjet injector struts as both ejector nozzles to entrain air at low speeds and later as the rocket combustion chambers for the ultimate ascent.
Whilst the scramjet engine unit is thermodynamically simple in conception, in engineering practice it's the most complex and technically demanding of all engine concepts mentioned in this paper. To make things worse many studies like the recent ESA 'Winged Launcher Concept' study have failed to show a good payload for a scramjet run SSTO because the important propulsive characteristics of scramjets are inadequately suited to the launcher role. The reduced specific thrust and high specific impulse of scramjets will favour a luxury cruise vehicle application flying at set Mach amount over long distances, especially since this would enable the elimination of the majority of the adjustable geometry.
Scramjet machines have a comparatively low specific thrust (nett thrust/air flow) due to the moderate combustor heat range rise and pressure proportion, and therefore an extremely large air mass move is required to give satisfactory vehicle thrust/weight ratio. However at frequent freestream dynamic brain the captured air mass circulation reduces for confirmed intake area as acceleration rises above Mach 1. Consequently the complete vehicle frontal area is needed to serve as an absorption at scramjet rates of speed and in the same way the exhaust movement must be re-expanded back into the initial streamtube in order to achieve an acceptable exhaust speed. However employing the automobile forebody and aftbody within the propulsion system has many cons:
In order to focus the intake distress system and make the right duct move areas over the complete Mach range, variable geometry intake/combustor and nozzle areas are required. The large variation in movement passage shape causes the adoption of an rectangular engine cross section with flat moving ramps in so doing incurring a severe charges in the pressure vessel mass. Also to increase the installed engine unit performance requires a high active pressure trajectory which in combo with the high Mach quantity imposes severe heating system rates on the airframe. Dynamic air conditioning of significant portions of the airframe will be necessary with further penalties in mass and difficulty.
Further downsides to the scramjet notion are visible in many areas. The nett thrust of a scramjet engine unit is very very sensitive to the intake, combustion and nozzle efficiencies due to the remarkably poor work proportion of the routine. Because the exhaust velocity is merely slightly higher than the incoming freestream velocity a small reduction in pressure recovery or combustion efficiency is likely to convert a small nett thrust into a little nett drag. This example might be tolerable if the theoretical methods (CFD codes) and executive knowledge were on a very sturdy footing with adequate correlation of theory with test. However the the truth is that the element efficiencies are determined by the specific physics of badly grasped areas like flow turbulence, shock wave/boundary layer relationships and boundary level change. To exacerbate this deficiency in the underlying physics existing surface test facilities cannot replicate the flowfield at bodily representative sizes, forcing the adoption of expensive air travel research vehicles to obtain the necessary data.
Scramjet development could only continue after an extended technology program and even then would probably be a risky and expensive job. In 1993 Effect Engines estimated a 130 tonne scramjet vehicle development program would cost $25B (at set prices) let's assume that this program proceeded matching to plan. This program would have included two X planes, one devoted to the subsonic handling and low supersonic program and the other an air decreased scramjet research vehicle to explore the Mach 5-15 regime.
In this section are grouped those engines that use turbocompressors to compress the airflow but without the aid of precoolers. The good thing about cycles that utilize onboard work copy to the air flow is they are capable of operation from sea level static conditions. This has important performance advantages over engines employing exclusively ram compression and also allows a cheaper development program since the mechanical reliability can be acquired in relatively inexpensive available air ground test facilities.
Turbojets (Fig. 1) exhibit a very quick thrust decay above about Mach 3 because of the effects of the rising compressor inlet temperatures forcing a decrease in both flow and pressure percentage. Compressors must be run within a stable part of the quality bounded by the surge and choke restrictions. Furthermore structural factors impose an top outlet temperatures and spool acceleration limit. As inlet heat increases (whilst operating at regular WЛ†T/P and N/Л†T) the spool quickness and/or outlet temp limit is quickly approached. Either way it's important to throttle the engine unit by moving down the jogging line, in the process reducing both stream and pressure proportion. The consequent decrease in nozzle pressure proportion and mass circulation results in an instant damage in nett thrust.
However at Mach 3 the automobile has received an insufficient boost to replace the mass charges of the airbreathing engine. Therefore each one of these cycles tend to be proposed together with a subsonic combustion ramjet mode to raised Mach statistics. The turbojet would be isolated from the hot airflow in ramjet mode by blocker gates which allow the airstream to flow around the central engine with small pressure reduction. The ramjet function provides sensible specific thrust to around Mach 6-7 of which point change to rocket propulsion is effected.
Despite the ramjet expansion to the Mach quantity range the performance of the systems is poor due primarily to their low thrust/weight percentage. An uninstalled turbojet has a thrust/weight ratio of around 10. However this falls to 5 or less when the consumption and nozzle systems are added which compares terribly with a H2/O2 rocket of 60+.
The turborocket (Fig. 2) cycles symbolize an attempt to improve on the low thrust/weight of the turbojet and improve the useful Mach quantity range. The genuine turborocket consists of a minimal pressure ratio admirer driven by an completely independent turbine employing H2/O2 combustion products. Because of the individual turbine working substance the coordinating problems of the turbojet are eased since the compressor can in theory be operated anywhere on its attribute. By manufacturing the compressor components in a suitable high temperature materials (such as reinforced ceramic) it is possible to get rid of the ramjet bypass duct and operate the engine unit to Mach 5-6 whilst being within outlet temperature and spool acceleration limits. Used this involves operating at reduced nondimensional velocity N/Л†T and therefore pressure ratio. As a result to avoid choking the compressor store guide vanes a low pressure ratio compressor is determined (often only 2 stages) which allows operation over the wider move range. The turborocket is substantially lighter when compared to a turbojet. Nevertheless the low pattern pressure ratio reduces the specific thrust at low Mach amounts and in conjunction with the preburner liquid air flow ends up with an unhealthy specific impulse compared to the turbojet.
This routine is a version of the turborocket whereby the turbine working fluid is replaced by ruthless regeneratively warmed hydrogen warmed in a heat exchanger positioned in the exhaust duct (Fig. 3). Due to heat exchanger material temperature restrictions the combustion process is generally put into two phases (upstream and downstream of the ma-
Fig. 1 Turbo-ramjet Engine motor (with built in rocket engine unit).
Fig. 2 Turborocket.
Fig. 3 Turbo-expander engine motor.
trix) and the turbine access temperatures is quite low at around 950K. This variant displays a average improvement in specific impulse compared with the clean turborocket because of the elimination of the water oxygen flow. Financial firms achieved at the expense of additional pressure loss in the air ducting and the mass charges of the heat exchanger.
Unfortunately none of them of the above engines display any performance improvement more than a pure rocket method of the SSTO launcher problem, regardless of the wide versions in core engine cycle and equipment. That is for the easy reason that the primary engine masses are swamped by the much larger public of the absorption and nozzle systems which have a tendency to outweigh the good thing about increased specific impulse.
Due to the relatively low pressure percentage ramjet modes of the engines, it is essential to offer an efficient ruthless recovery varying geometry intake and a changing geometry exhaust nozzle. The necessity for high pressure recovery pushes the adoption of 2 dimensional geometry for the consumption system because of the requirement to focus multiple oblique shockwaves over a wide mach quantity range. This results in a very serious mass penalty due to the inefficient pressure vessel mix section and the actually large and complicated moving ramp assemblage using its high actuation lots. Similarly the exhaust nozzle geometry must manage to a broad area ratio variance in order to cope with the widely differing flow conditions (WЛ†T/P and pressure proportion) between transonic and high Mach quantity flight. An additional complication emerges because of the requirement to assimilate the rocket engine motor necessary for the later ascent in to the airbreathing engine nozzle. This avoids the prohibitive bottom drag penalty that could result from another 'lifeless' nozzle system as the automobile attempted to accelerate through transonic.
Liquid Air Cycle Motors were first proposed by Marquardt in the early 1960's. The simple LACE engine exploits the low heat and high specific heating of liquid hydrogen to be able to liquify the captured airstream in a specially designed condenser (Fig. 4). Pursuing liquifaction the air is relatively easily pumped up to such high stresses that it could be fed into a typical rocket combustion chamber. The main advantage of this process would be that the airbreathing and rocket propulsion systems can be combined with only a single nozzle required for both modes. This brings about a mass saving and a compact installation with productive platform area utilisation. Also the engine unit is in process capable of operation from sea level static conditions up to perhaps Mach 6-7.
Liquid Air Turbopump Fig. 4 Water Air Cycle Engine unit (LACE).
The main disadvantage of the LACE engine however is that the fuel utilization is very high (in comparison to other airbreathing motors) with a specific impulse of only about 800 secs. Condensing the air flow necessitates the removal of the latent high temperature of vaporisation under isothermal conditions. However the hydrogen coolant is within a supercritical talk about following compression in the turbopump and absorbs the heat insert with an accompanying increase in temps. Consequently a temps 'pinch point' occurs within the condenser at around 80K and can only be handled by increasing the hydrogen stream to many times stoichiometric. Air pressure within the condenser impacts the latent warmth of vaporisation and the liquifaction temp and consequently has a solid effect on the gas/air percentage. However at sea level static conditions of around 1 bar the minimum fuel/air ratio required is about 0. 35 (ie: 12 times greater than the stoichiometric ratio of 0. 029) assuming that the hydrogen have been compressed to 200 club. Increasing mid-air pressure or minimizing the hydrogen pump delivery pressure (and temps) could decrease the gas/ air ratio to perhaps 0. 2 but nevertheless the fuel stream remains very high. At high Mach volumes the fuel move may need to be increased further, due to heating exchanger metal heat limitations (exacerbated by hydrogen embrittlement limiting the decision of pipe materials). To lessen the fuel flow it is sometimes proposed to hire slush hydrogen and recirculate some of the coolant move back into the tankage. Nevertheless the handling of slush hydrogen poses difficult specialized and functional problems.
From a technology standpoint the key challenges of the easy LACE engine are the need to prevent clogging of the condenser by frozen skin tightening and, argon and water vapour. Also the ability of the condenser to handle a changing 'g' vector and of creating a scavenge pump to use with a very low NPSH inlet. Nevertheless performance studies of SSTO's prepared with LACE motors show no performance benefits because of the inadequate specific impulse in airbreathing method despite the fair thrust/weight ratio and Mach number capability.
The Air Collection Engine unit (ACE) is a far more complex variant of the LACE engine when a liquid oxygen separator is incorporated following the air liquifier. The intent is to takeoff with the primary liquid oxygen tanks bare and load them through the airbreathing ascent therefore possibly lowering the undercarriage mass and installed thrust level. The ACE primary is often proposed for parallel operation with a ramjet main propulsion system. On this version the hydrogen energy movement would condense a quantity of air that the air would be segregated before going into the ramjet combustion chamber at a near stoichiometric mix percentage. The liquid nitrogen from the separator could perform various chilling duties before being given back into the ramjet air flow to recover the momentum pull.
The oxygen separator will be a sophisticated and heavy item because the physical properties of liquid air and nitrogen are very similar. However putting away the anatomist details, the essential thermodynamics of the ACE main are wholly unsuited to an SSTO launcher. Since a petrol/air mixture proportion of around 0. 2 is required to liquify the environment and since oxygen is 23. 1% of the air flow it is apparent that a about identical mass of hydrogen is required to liquify a given mass of oxygen. Therefore there is absolutely no cutting down in the takeoff propellant loading and the truth is a severe structure mass penalty because of the increased fuselage size needed to contain the low density water hydrogen.
This last course of motors is specifically designed for the SSTO propulsion role and combines some of the best features of the prior types whilst simultaneously conquering their faults. The first engine of this type was the RB545 powerplant suitable for the HOTOL spaceplane and formerly devised by Alan Relationship in 1982. Building on the knowledge gained throughout the HOTOL job the thermodynamics of this first engine unit were refined during 198990 leading to the SABRE powerplant designed for SKYLON.
The global specs for this type of engine results from an examination of the necessary propulsion characteristics for a successful SSTO spaceplane:
From the aforementioned list it is clear that the LACE engine is close to meeting the requirements aside from its high gasoline consumption. Therefore the RB545 engine unit resulted as an development from the Ribbons cycle in order to improve its specific impulse.
The excessive energy circulation of the Ribbons engine is completely because of the quantity of coolant necessary to result the condensation process. The work capacity of the hot ruthless gaseous hydrogen stream that emerges from the precooler/condenser remains basically unexploited since the power demands of the liquid air turbopump are little. The RB545 circuit was born out of acceptance that a better split between the cooling and work demands of the pattern could be performed by preventing the air liquifaction process entirely. By terminating the cooling down process near to but not below the vapour boundary (Л†80K) the 'pinch point' was avoided with a very large saving in the mandatory coolant movement, whilst still departing sufficient hydrogen to drive a high pressure ratio turbocompressor sufficiently powerful to compress the air flow up to typical rocket combustion chamber pressures. With this routine the most effective compressor inlet temperatures that minimises the total hydrogen flow is actually on the vapour boundary since the compressor work needs are higher than the air conditioning requirements. Nevertheless by deep precooling of the inbound airstream the compressor work demand is greatly reduced and increased compressor outlet temperature ranges are avoided particularly at high Mach statistics. Also unlike a straightforward turbojet the engine unit does not have problems with a reduction in gross thrust with increasing Mach quantity because the precooler 'irons out' the intake recovery air temperature variance allowing the compressor to operate with a nearly constant inlet temperatures.
In practice the RB545 used the high pressure hydrogen delivery from the hydrogen turbopump to cool the airstream immediately, following that your hydrogen stream divide. Approximately onethird handed to the key combustion chamber via the preburner whilst the remaining two-thirds was broadened through the turbocompressor turbine prior to exhaust. This circuit reduced the gasoline/air ratio to around 0. 1. However this was eroded at high Mach statistics due to precooler steel temperature limitations caused by hydrogen embrittlement. Apart from the better specific impulse of this cycle the majority of the technology problems of the Ribbons engine are avoided (eg: two phase temperature exchangers and liquid air handling). The look of the turbomachinery and heat exchanger areas are relatively classic although there remains the situation of avoiding atmospheric moisture content clogging the precooler with frost. Somewhat surprisingly the total engine mass is not a greater than the LACE engine because the addition of the turbocompressor is approximately balanced by the elimination of air condenser and turbopump. The turbocompressor is a lot lighter than an equivalent compressor sketching ambient air because the low air inlet temperature reduces the physical size of the unit because of the higher air density, and also reduces the rotational speeds (and hence inertial loading) due to the lower speed of audio. Also the low air delivery heat range allows light alloy or amalgamated construction for almost all of the compressor which combined with the previous factors reduces the mass to roughly one-quarter that of an ambient machine.
The SABRE engine motor (Figs. 5 and 6) is a more complex variant of the initial cycle in which a lower fuel flow has been achieved at the expense of a tiny mass penalty. In such a engine a Brayton circuit helium loop has been interposed between your 'hot' airstream and the 'chilly' hydrogen stream. The task outcome of the helium loop supplies the power to drive mid-air compressor. Employing helium as the working substance permits superior temperature resisting alloys for the precooler matrix and also results in more optimally matched turbine stages. The superior thermodynamics of the SABRE engine cause an air/fuel ratio of about 0. 08 which produces a useful
Fig. 5 SABRE engine motor thermodynamic pattern.
Fig. 6 SABRE vertical combination section.
saving in hydrogen ingestion during the airbreathing ascent. Frost control is effected in a more efficient manner on SABRE which also has a favourable impact on the vehicles' payload potential.
A very important factor in the success of the engines is the fact that because of the high internal pressure percentage of the central engine it is not necessary to equip the automobile with an extremely efficient absorption system. The SABRE engine unit operates over the complete trajectory with an inlet pressure of only around 1. 3 pub, which allows maximum chamber pressure to be achieved with minimal versions in the turbomachinery operating point. Subsequently a 2 distress intake (one oblique and one normal distress) can meet the engine demands, permitting the complicated and heavy 2-dimensional intakes typical of low pressure ratio motors to be dispensed with in favour of a straightforward axisymmetric inlet with a translating centrebody. The simplification of the absorption mechanism and increases in size from an axisymmetric framework bring about an intake mass keeping of around 80% compared with a higher pressure recovery 2D intake. Also, since forebody precompression is needless, the intake assemblage can be removed from within the vehicle dramatically improving the design flexibility to solve cut and airframe layout problems.
At rocket transition air inlet is shut and the turbocompressor is run-down whilst concurrently the liquid air turbopump is run-up. The preburner heat range is low in rocket setting reflecting the reduced power demand of the liquid oxygen turbopump. The engine includes a liquid air cooled main combustion chamber since this allows the same oxidiser injectors to be used in both methods.
The installed specific impulse and thrust/weight proportion of the SABRE engine are shown in Figs. 7 and 8 with the other engine motor prospects shown for comparison. It is important to note that the individuals have been evaluated using broadly extant materials and aerothermodynamic technology. These characters show that the SABRE engine motor achieves a specific impulse equivalent with turborockets whilst concurrently attaining installed thrust/ weight ratios a lot like LACE engines. It really is this com-
Fig. 7 Installed thrust/weight proportion of the blended SSTO propulsion assembly in airbreathing setting.
Fig. 8 Installed specific impulse of SSTO propulsion systems
bination of modest specific impulse with low installed weight which makes precooled hybrid machines uniquely well suited for SSTO kick off vehicles.
The program of the SABRE engine motor is described in more detail in  which also protects the design of
a suited airframe (SKYLON) which properly harnesses the entire potential and unique characteristics of the engine type. The final SABRE/SKYLON mixture is capable of placing a 12 tonne payload into an equatorial low Globe orbit at a gross takeoff mass of 275 tonnes (payload portion 4. 36%).
R. Varvill and A. Connection, "The SKYLON Spaceplane", Paper
IAA 95-V3. 07, offered at the 45th IAF Congress, Oslo, 1995.
B. R. A. Melts away, "HOTOL space transportation for the twenty first century", Proceedings of the Institute of Mechanical Technical engineers, 204, pp. 101-110, 1990.
(Received 14 November 2002)
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