Article

The Skylon spaceplane: Progress to realisation

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  • Reactionengines
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Abstract

The Skylon spaceplane will enable single stage to orbit delivery of payloads with aircraft like operations. The key to realising this goal is a combined cycle engine that can operate both in airbreathing and pure rocket modes. To achieve this new low mass structure concepts and several new engine technologies need to be proven. An extensive program of technology development has addressed these issues with very positive results. This now allows the project to proceed to the final concept proving stage before full development commences.

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... The NASP programme was cancelled because its design demanded an airbreathing scramjet engine to operate at Mach g o acceleration due to earth's gravity, 9.81 m/s 2 [3,4] and the Hyperplane/Avatar as they meet all the nine attributes of true spaceplanes. Skylon C2 (345 tonnes take-off weight/15 tonnes payload weight) with 25% HFF is a larger version of Skylon C1 (275/12), with a take-off capable undercarriage; while the 198-tonne Hotol was the original version that had a specially designed detachable trolley for fully fuelled take-off and light-weight undercarriage for landing. ...
... Among SSTO reusable spaceplanes studied, the Skylon C2 assumes a life-cycle of 200 flights and a specific launch cost of $1300 per kg in LEO based on fixed and variable costs spread over vehicle life cycle of 200 flights [3,17].In the case of the Saenger TSTO, the first stage is designed as a passenger air-craft, and the second stage alone is a reusable spaceplane. The passenger aircraft will experience some 20,000 take-off and landings and 50,000 hours flight time whereas the second stage space launch vehicle will be exposed to only 150-200 launches and 500-1000 hours flight time [19] with a specific launch cost of $1867/kg in LEO at 1985 prices (or $3063/kg current prices at 2% inflation rate over 25 years). ...
... Remarkable progress in design, development and production of stainless steel and inconel cryogenic heat exchangers for air cooling up to a speed of Mach 5, temperatures up to 1300K i.e. the precooler, has been accomplished in the UK through the Sabre engine and Skylon vehicle concept [3,4,27,28]. Major problems like frosting of the heat exchanger during ascent flight in the atmosphere have also been successfully addressed. ...
Conference Paper
Full-text available
Mankinds' efforts throughout the latter half of the 20th.Century to develop reusable spaceplanes have not succeeded because their design requirements called for large vehicle size and weight, entailing high development cost and unacceptable technical risk. An aerobic, folly reusable SSTO Spaceplane design permits technology development to commence with a small size vehicle having a take-off weight as low as 25 tonnes delivering a payload of one tonne in low-earth orbit. Aerobic vehicle payload fraction is influenced by take-off weight, choice of propulsion system, powered aerodynamic maneuvers for Lox collection, aerodynamic-configuration and airframe structural design considerations. Extensive flight performance optimization studies for these parameters were done. Optimizing the modified Kuchmann factor minimized the size and weight of the aerobic vehicle. Configuration design of "Avatar" An Aerobic SSTO Aerospaceplane, for which International Design Patent Registration Application has been filed, is presented. Design limitations and design margins available are discussed. Critical areas in technology development and their sources are identified. The need for a cooperative global approach to Spaceplane design and development is established. A long-term perspective plan for development and deployment of a Space Solar Power System as a global energy resource using Aerobic Spaceplanes is the proposed space mission. © 2000 by Government of India. Published by the American Institute of Aeronautics and Astronautics, Inc.
... Saenger, Hope, Hermes, and Buran dropped out of consideration possibly due to budgetary constraints and political decisions and it may be pointed out that their propulsion design concepts ended up (in retrospect) in relatively low HFF (8-15%) at take-off. Thus, the surviving concepts are the Hotol [2]/Skylon [3,4] and the Hyperplane/Avatar as they meet all the nine attributes of true spaceplanes. Skylon C2 (345 tonnes take-off weight/15 tonnes payload weight) with 25% HFF is a larger version of Skylon C1 (275/12), with a take-off capable undercarriage; while the 198-tonne Hotol was the original version that had a specially designed detachable trolley for fully fuelled take-off and light-weight undercarriage for landing. ...
... However, the life-cycle of spaceplanes is both a technical factor as well as an economic factor, unlike fleet size, flight rate, development investment amortization etc. Hence the life-cycle factor could be considered in the estimations of relative technical cost. Among SSTO reusable spaceplanes studied, the Skylon C2 assumes a life-cycle of 200 flights and a specific launch cost of $1300 per kg in LEO based on fixed and variable costs spread over vehicle life cycle of 200 flights [3,17].In the case of the Saenger TSTO, the first stage is designed as a passenger air-craft, and the second stage alone is a reusable spaceplane. The passenger aircraft will experience some 20,000 take-off and landings and 50,000 hours flight time whereas the second stage space launch vehicle will be exposed to only 150-200 launches and 500-1000 hours flight time [19] with a specific launch cost of $1867/kg in LEO at 1985 prices (or $3063/kg current prices at 2% inflation rate over 25 years). ...
... Remarkable progress in design, development and production of stainless steel and inconel cryogenic heat exchangers for air cooling up to a speed of Mach 5, temperatures up to 1300K i.e. the precooler, has been accomplished in the UK through the Sabre engine and Skylon vehicle concept [3,4,27,28]. Major problems like frosting of the heat exchanger during ascent flight in the atmosphere have also been successfully addressed. ...
Article
Full-text available
A new concept for flight to orbit is described in this paper. It involves mass addition to an ascending air-breathing, hypersonic lifting vehicle. General laws for flight to orbit with mass addition are developed, and it is shown that payload capabilities one order of magnitude higher than even the most advanced rocket launchers is feasible. Detailed concept definition and design studies are presented for this new hydrogen fueled, horizontal take off, fully reusable single stage hypersonic vehicle, called HYPERPLANE. The studies include results from over 5000 trajectory optimization runs on an interactive mission analysis computer model; an analysis of various propulsion options; and design and performance optimization of an air collection system for mass addition in flight.
... Hypersonic vehicles are the next generational aircrafts in or cross the atmosphere, using the air-breathing [1], combined cycle [2,3] or pre-cooled engines [4]. To achieve a long endurance hypersonic flight, a high-power electricity generation system is essential for supplying the fuel feeding, environment control and radar system [5]. ...
... Generally, there are two kinds of fuel suitable for hypersonic vehicles, namely the cryogenic fuel (such as liquid hydrogen) and hydrocarbon fuel at normal temperature (such as kerosene). Liquid hydrogen has huge physical heat sink and high heat value per unit mass, making it an ideal fuel for the pre-cooled engines [4] and the air-breathing propulsion at high Mach numbers (Ma>10) [17]. But the low-temperature storage of hydrogen is technologically difficult on the leakage and the size of fuel tank, because of the small molecule and the low heat value per unit volume. ...
... To utilize the cryogenic exergy of liquid hydrogen, the CBC working fluid is chosen as helium. Helium cannot be liquidated by liquid hydrogen, so that the working fluid keeps gaseous state in the loop [4]. For helium CBC system, the interstage cooling is an effective technology to improve cycle performance [18]. ...
Article
Electricity generation by means of closed-Brayton-cycle (CBC) systems on hypersonic vehicles is strictly limited by the finite cold source, namely the onboard fuel. The influences of finite cold source on electricity generation, and the performance comparison among the CBC systems with different fuels, are worth investigating. In this article, the main factors and their effects on the electric power of CBC systems with finite cold source are analyzed. A simple recuperated CBC model has been established to evaluate the performances of power generation. Results indicate that the available temperature zones of cold source have great influence on the efficiency and power of CBC systems. The cryogenic fuels have greater potential of power generation than the hydrocarbon fuels at normal temperature, because of the advantages on the thermal efficiency (45% vs. 27%), effective enthalpy difference (5.67 MJ/kg vs. 1.06 MJ/kg) and effectiveness of primary cooler (∼0.9 vs. ∼0.85). Specifically, a higher electric power fraction (9.46%) can be achieved by the CBC system cooled by liquid hydrogen than kerosene (2.61%). This research provides the performance boundaries of onboard closed power generation systems in the view of finite cold source.
... Saenger, Hope, Hermes, and Buran dropped out of consideration possibly due to budgetary constraints and political decisions and it may be pointed out that their propulsion design concepts ended up (in retrospect) in relatively low HFF (8-15%) at take-off. Thus, the surviving concepts are the Hotol [2]/Skylon [3,4] and the Hyperplane/Avatar as they meet all the nine attributes of true spaceplanes. Skylon C2 (345 tonnes take-off weight/15 tonnes payload weight) with 25% HFF is a larger version of Skylon C1 (275/12), with a take-off capable undercarriage; while the 198-tonne Hotol was the original version that had a specially designed detachable trolley for fully fuelled take-off and light-weight undercarriage for landing. ...
... However, the life-cycle of spaceplanes is both a technical factor as well as an economic factor, unlike fleet size, flight rate, development investment amortization etc. Hence the life-cycle factor could be considered in the estimations of relative technical cost. Among SSTO reusable spaceplanes studied, the Skylon C2 assumes a life-cycle of 200 flights and a specific launch cost of $1300 per kg in LEO based on fixed and variable costs spread over vehicle life cycle of 200 flights [3,17].In the case of the Saenger TSTO, the first stage is designed as a passenger air-craft, and the second stage alone is a reusable spaceplane. The passenger aircraft will experience some 20,000 take-off and landings and 50,000 hours flight time whereas the second stage space launch vehicle will be exposed to only 150-200 launches and 500-1000 hours flight time [19] with a specific launch cost of $1867/kg in LEO at 1985 prices (or $3063/kg current prices at 2% inflation rate over 25 years). ...
... Remarkable progress in design, development and production of stainless steel and inconel cryogenic heat exchangers for air cooling up to a speed of Mach 5, temperatures up to 1300K i.e. the precooler, has been accomplished in the UK through the Sabre engine and Skylon vehicle concept [3,4,27,28]. Major problems like frosting of the heat exchanger during ascent flight in the atmosphere have also been successfully addressed. ...
Article
Full-text available
A new concept for flight to orbit which involves mass addition to an ascending, air-breathing, hypersonic lifting vehicle is described. Studies performed on this hyperplane are detailed and results of trajectory optimization runs carried out on an interactive mission analysis computer model are presented. Various propulsion options are analyzed and a design and performance optimization is carried out on an air collection system for mass addition in flight.
... Therefore, various Thermal Protection Systems (TPS) are in use today which protect the payload from the shocked hot gas. In addition to the TPS needs of hypersonic spacecraft in use today, envisaged re-usable hypersonic passenger vehicles, such as the Skylon [2], will require protection from high heat loads [3]. Therefore, an interest in further research on re-usable and high performing thermal protection systems exists, where transpiration cooling is considered as one of the most promising candidates [4]. ...
... The intermediate results are showcased in Fig. 5 for UHTC-3 (post-sand). The normalised differential pressure is plotted against the superficial input velocity including the uncertainties of the experimental data in form of error bars in Fig. 5a; the solid line here, which represents the Darcy-Forchheimer equation, is obtained by using the K D and K F values in Eq. (2). The corresponding mass flow rate through an area of A = 1.911 × 10 −4 m 2 on the porous sample, is presented in Fig. 5b. ...
Conference Paper
This paper experimentally examines the internal and external flow characteristics of porous zirconium diboride (ZrB2), an Ultra-High-Temperature-Ceramic (UHTC) and a potential candidate for transpiration cooling of hypersonic vehicles. This is performed for both partially sintered material and fully densified material with cast features. The Darcy and Forchheimer permeability coefficients of these samples are determined using an ISO standard test rig. The outflow of the transpiring porous samples is investigated where no hypersonic cross-flow is involved using hot-wire anemometry and focused Schlieren visualisation. The velocity maps obtained from the hot-wire data show significant non-uniformities across the UHTC’s outflow region, both at low and high differential pressures. The focused Schlieren using carbon dioxide as the injected gas reveals unsteady structures at high differential pressures as the outflowing gas interacts with the surrounding air.
... Because SSTO RLVs were not feasible with rocket engines, the basic idea for the SSTO HOTOL vehicle (1982)(1983)(1984)(1985)(1986)(1987)(1988)(1989) was to augment rocket engine performance using the atmosphere instead of starting with an air-breathing engine, and then to find some way to add a rocket engine for the latter part of the flight [34]. This approach was used to design the HOTOL propulsion system, the Rolls Royce RB545 engine, which was a combined-cycle precooled engine [35]. ...
... The conceptual, unpiloted, reusable Skylon aerospace plane (1989 to present), an SSTO RLV, which can be used as a reusable first stage of a TSTO launch system ( [36] pp. [35][36][37][38][39], evolved from HOTOL. For this concept, various propulsion concepts, hydrogen/oxygen rockets, scramjets, turbojets, turborockets, and liquid air cycle engines, were considered [37]. ...
Article
The Skylon concept incorporates the highly innovative synergetic air-breathing rocket engine concept that has the potential to revolutionize the mode of propulsion for transportation of medium-weight payloads to low Earth orbits. An independent partial assessment is provided of the technical viability of the Skylon concept. Pressure lift and drag coefficients derived from Euler simulations for unpowered flight compare very well and fairly well, respectively, with those from engineering methods. The engineering-method coefficients for powered flight are increasingly less acceptable as the freestream Mach number is increased beyond 8.5 because these methods did not account for the increasing favorable (in terms of pressure forces) effect of underexpanded rocket engine plumes on the aft fuselage. At Mach numbers greater than 8.5, the thermal environment around the aft fuselage is a known unknown: a potential design and/or performance risk issue. The adverse effects of shock waves on the aft fuselage and plume-induced flow separation are other potential risks. A preliminary design of Skylon requires the judicious use of a combination of engineering methods, advanced methods based on required physics or analytical fidelity, test data, and independent assessments. The demonstration of a synergetic air-breathing rocket-engine-powered experimental aerospace plane calls for the second revival of the Aerospace Plane Program.
... An air-breathing horizontal take-off launch vehicle is seen as a revolutionary step in space transportation, enabling dramatically safer, more affordable and reliable, re-usable launch system with enhanced flexibility in operations [5][6][7]. One enabler of the revolutionary step is that oxygen is taken from the atmosphere, rather than being extra mass carried on the vehicle, as is the case for rockets. ...
... In a recent program the NASA X-43 set out to develop technology for the first stage of a two stage to orbit vehicle and was successful, operating at Mach 9.68 [5]. Another launch vehicle concept is the Skylon single stage to orbit vehicle designed around an air-breathing and rocket powered Sabre engine by Reaction Engines [6]. ...
Article
The design and optimization of hypersonic aircraft is severely impacted by the high temperatures encountered during flight as they can lead to high thermal stresses and a significant reduction in material strength and stiffness. This reduction in rigidity of the structure requires innovative structural concepts and a stronger focus on aeroelastic deformations in the early design and optimisation of the aircraft structure. This imposes the need for a closer coupling of the aerodynamic and structural design tools than is current practice. The paper presents the development of a multi-disciplinary, closely coupled optimisation suite for hypersonic aircraft. An overview of the setup and structure of the optimization suite is given and the integration between the Tranair solver, used to determine the aerodynamic loads and temperatures, and MSC/NASTRAN, used for the structural sizing and design, will be given.
... Furthermore, the development status of such engines is reported in the literature. However, it is mainly focused on the progress of the pre-cooler and other key technologies, as well as the infrastructure of the RLV to be equipped with these engines (9,10,(17)(18)(19)(20)(21)(22)(23) . ...
Article
Full-text available
This paper presents a performance analysis on a novel engine concept, currently under development, in order to achieve hybrid air-breathing rocket technology. A component-level approach has been developed to simulate the performance of the engine at Mach 5, and the thermodynamic interaction of the different working fluids has been analysed. The bypass ramjet duct has also been included in the model. This facilitates the improved evaluation of performance parameters. The impact of ram drag induced by the intake of the engine has also been demonstrated. The whole model is introduced into a multi-platform application for aeroengine simulation to make it accessible to the interested reader. Results show that the bypass duct modelling increases the overall efficiency by approximately 7%. The model calculates the specific impulse at approximately 1800 seconds, which is 4 times higher than any chemical rocket.
... The hypersonic precooled engine is a potential propulsion system for hypersonic flight and space access [1,2], which can efficiently operate over a wide range of Mach numbers, with low emissions, high specific impulse as well as high thrust-to-weight ratios [3,4,5]. The precooler with high performance and robustness is a key component of this highperformance engine, which quenches the intake air temperature under high Mach number flight conditions [6][7][8]. ...
Article
The precooler with high power-to-weight ratios and robustness is a core component of hypersonic precooled engines to operate efficiently at the full speed range. In order to figure out the feasibility of precooler technology, an annular microtube-typed precooler was designed and manufactured. A high temperature test platform consisting of an extreme thermal environment system and a closed S-Helium loop system was built to experimentally verify the operating performance and reliability of this novel precooler. Moreover, a parameter matching method was proposed for stably operating the test platform. On these basis, an 1800-seconds longtime precooler performance test simulating Mach4 heat conditions and constantly supplying S-Helium at 8MPa was successfully conducted. The test results are in a great agreement with the model results and the maximum error of the heat transfer rate is within ±10%. The manufactured microtube-typed precooler exhibits its long-term resistance to harsh operating environment (at temperature of nearly 1000K and pressure of over 8MPa) and maintains its structural integrity after thermal cycles with an acceptable leakage rate of below 3vol%/h. The airflow temperatures were quenched in excess of 600K in the test which demonstrates the precooler’s ability to cool airflow at speeds significantly beyond the limit of any jet-engine powered aircraft. In addition, the measured power/weight ratio is up to 100kW/kg, which is in the lead in precooler technology to the authors’ knowledge.
... A vital pre-cooler technology relates to preventing the matrix clogging with frost due to condensation of water vapour in the engine airflow. REL have developed an efficient solution which was developed in a purposebuilt frost control wind tunnel in the early 2000s [4]. ...
Article
Full-text available
SKYLON is a reusable single stage to orbit spaceplane that can take off from a runway reach a 300 km altitude low earth orbit with a payload of 15 tonnes and then return to Earth for a runway landing. The unique feature of SKYLON that enables it to achieve this objective is the Synergistic Air-Breathing Rocket Engine (SABRE) which has both air breathing and pure rocket modes. The SKYLON development programme has concentrated on the SABRE engines and the component level technology development programme was completed in 2013 with the successful demonstration of a complete pre-cooler system using flight representative modules. The technology readiness has reached the point when the next phase of the development programme has begun. This £250 million programme will demonstrate the technologies in a system complex and take the design of the flight SABRE to CDR. The SKYLON airframe has also been subject of a slower paced programme including a series of technology development projects, mostly centred on the structure and thermal protection system, and a revision of the system design to incorporate the results of both the airframe and engine programmes.
... 29 This approach was used to design the HOTOL propulsion system−the Rolls Royce RB545 engine−which was a combined cycle pre-cooled engine. 30 The conceptual, unpiloted, reusable Skylon aerospace plane−an SSTO RLV, which can be used as a reusable first stage of a TSTO launch system 31 −(1989-present) evolved from HOTOL. For this concept, various propulsion concepts-hydrogen/oxygen rockets, scramjets, turbojets, turborockets, and Liquid Air Cycle Engines-were considered. ...
Conference Paper
Full-text available
An independent partial assessment is provided of the technical viability of the Skylon aerospace plane concept, developed by Reaction Engines Limited (REL). The objectives are to verify REL's engineering estimates of airframe aerodynamics during powered flight and to assess the impact of Synergetic Air-Breathing Rocket Engine (SABRE) plumes on the aft fuselage. Pressure lift and drag coefficients derived from simulations conducted with Euler equations for unpowered flight compare very well with those REL computed with engineering methods. The REL coefficients for powered flight are increasingly less acceptable as the freestream Mach number is increased beyond 8.5, because the engineering estimates did not account for the increasing favorable (in terms of drag and lift coefficients) effect of under-expanded rocket engine plumes on the aft fuselage. At Mach numbers greater than 8.5, the thermal environment around the aft fuselage is a known unknown−a potential design and/or performance risk issue. The adverse effects of shock waves on the aft fuselage and plume-induced flow separation are other potential risks. The development of an operational reusable launcher from the Skylon concept necessitates the judicious use of a combination of engineering methods, advanced methods based on required physics or analytical fidelity, test data, and independent assessments. Nomenclature Symbols A t = nozzle throat area A e = nozzle exit area C L = pressure lift coefficient C D = pressure drag coefficient C z = pressure moment around z-axis F x = pressure force in x-direction, the direction from nose to tail of Skylon F y = pressure force in y-direction, the vertical direction F z = pressure force in z-direction, the span-wise direction h = altitude J = objective function M ∞ = freestream Mach number M j = Jet Mach number at nozzle exit m a = airflow rate m f = fuel flow rate m e = mass flow rate at nozzle exit p ∞ = freestream pressure P 1t = total pressure at the exit of combustion chamber P e = static pressure at nozzle exit T 1t = total temperature at the exit of combustion chamber T e = static temperature at nozzle exit T rec = freestream recovery temperature T tot = freestream total temperature V e = velocity at nozzle exit
... The next generation of vehicles will be used to carry either payload or crews into space, and by emphasizing full re-usability in their design and employing an airline-like approach, where the cost of acquisition is amortized over repeated flights, these vehicles promise to dramatically reduce the cost per kilogram of access to space. An example of such a vehicle is the Skylon space plane that is currently under development within the UK by Reaction Engines Ltd. 1,2 The overarching aim of the research described in this paper is to develop a model-based software tool that will aid in the preliminary design of the next generation of space-access vehicles. Within this paper, an ascent trajectory, starting in the low supersonic regime and extending through to hypersonic velocities, is designed for a representative space plane concept by optimizing a control law that alters the angle of attack of the vehicle and the throttle setting of its engines to vary the flight time along its trajectory into orbit. ...
Conference Paper
Full-text available
This paper addresses the design of ascent trajectories for a hybrid-engine, high performance, unmanned, single-stage-to-orbit vehicle for payload deployment into low Earth orbit. A hybrid optimisation technique that couples a population-based, stochastic algorithm with a deterministic, gradient-based technique is used to maximize the final vehicle mass in low Earth orbit after accounting for operational constraints on the dynamic pressure, Mach number and maximum axial and normal accelerations. The control search space is first explored by the population-based algorithm, which uses a single shooting method to evaluate the performance of candidate solutions. The resultant optimal control law and corresponding trajectory are then further refined by a direct collocation method based on finite elements in time. Two distinct operational phases, one using an air-breathing propulsion mode and the second using rocket propulsion, are considered. The presence of uncertainties in the atmospheric and vehicle aerodynamic models are considered in order to quantify their effect on the performance of the vehicle. Firstly, the deterministic optimal control law is re-integrated after introducing uncertainties into the models. The proximity of the final solutions to the target states are analysed statistically. A second analysis is then performed, aimed at determining the best performance of the vehicle when these uncertainties are included directly in the optimisation. The statistical analysis of the results obtained are summarized by an expectancy curve which represents the probable vehicle performance as a function of the uncertain system parameters. This analysis can be used during the preliminary phase of design to yield valuable insights into the robustness of the performance of the vehicle to uncertainties in the specification of its parameters. © 2012 by Richard E Brown, Edmondo Minisci, Christie Maddock, Ian Taylor, Fabrizio Pescetelli.
... The acceleration phase is an important part of flight for most hypersonic vehicles, but it is particularly important for the application of space access. The most well-known proposal for such a concept is the NASP [26], but other programs such as Skylon [27,28] and the Astrium spaceplane [29] are the subjects of current research. Other concepts have been developed with the backing of very interesting research but without such popular names or acronyms. ...
Article
The development and application of a first-principles-derived reduced-order model called MASIV (Michigan/AFRL Scramjet In Vehicle) for an airbreathing hypersonic vehicle is discussed. Several significant and previously unreported aspects of hypersonic flight are investigated. A fortunate coupling between increasing Mach number and decreasing angle of attack is shown to extend the range of operating conditions for a class of supersonic inlets. Detailed maps of isolator unstart and ram-to-scram transition are shown on the flight corridor map for the first time. In scram mode the airflow remains supersonic throughout the engine, while in ram mode there is a region of subsonic flow. Accurately predicting the transition between these two modes requires models for complex shock interactions, finite-rate chemistry, fuel-air mixing, pre-combustion shock trains, and thermal choking, which are incorporated into a unified framework here. Isolator unstart occurs when the pre-combustion shock train is longer than the isolator, which blocks airflow from entering the engine. Finally, cooptimization of the vehicle design and trajectory is discussed. An optimal control technique is introduced that greatly reduces the number of computations required to optimize the simulated trajectory.
... Some argue that short term implementation is both possible and desirable, with the correct design approach. 3 The reduction in cost of access to space such a system would provide should clear the way for a much more expansive utilisation of space, and development of space infrastructure. Such vehicles inevitably require propulsion systems that operate from launch to orbit, and are thus subject to the limitations discussed above. ...
Conference Paper
Full-text available
Replacing conventional bell nozzles with altitude compensating forms represents an attractive proposition for launch vehicle design, as both effciency across the altitude range and altitude performance. through the use of larger area ratios. may be increased. This paper compares the performance of two such compensating nozzle types, the Expansion Deflection and Dual Bell, through a series of cold flow tests conducted at nozzle pressure ratios encompassing the entirety of the compensating regime. Nozzle effciencies are shown to vary significantly across the pressure ratio range. Furthermore, relatively small changes in some design parameters associated with the ED nozzle type are shown to have an effect on overall performance. Despite this, the compensating behaviour of the type is shown to have similar potential to that of the Dual Bell. Combined with its other advantages of shorter length and potentially more adaptable design principles, it appears to show promise for application to future launchers. © 2011 by the American Institute of Aeronautics and Astronautics, Inc.
... SABRE (Synergistic Air-Breathing Rocket Engine) ,went by the name of a revolutionary pre cooling combined cycle engine, was developed by Reaction Engines Ltd., which can be used as a power plant of a low cost reusable single stage called Skylon [5] . The reason SABRE has such a superior performance is the use of ultra-light heat exchanger, using helium as a cooling medium, which have ability of cooling the income flow from 1000 ℃ down to -150 ℃ at a mass rate of 400kg / s air within a few cent seconds [6] , The maximum heat exchange power is up to 400MW. Compared to the design of JAXA, Sabre adopts a more compact design, with supercritical helium as a cooling medium, so that heat transfer per unit mass of precooler is greater. ...
Conference Paper
Full-text available
As a critical component of air-breathing hypersonic engine, an advanced compact heat exchanger is able to obtain high power to weight ratio with low pressure drop. Due to the special structure and coolant of the advanced compact heat exchanger, traditional design and performance-prediction methods are not appropriate for this kind of heat exchanger. In order to evaluate performance of compact heat exchangers reliably and validate the new design and performance-prediction approach, test facility and technology dedicated to experimental study for compact heat exchangers on flow and heat transfer are indispensable. Experimental apparatus, which consist of high temperature air tunnel and the coolant loop, is built out and a heat transfer experiment is conducted under representative working conditions for a certain compact heat exchanger in this paper. The preliminary experimental results show that the compact heat exchanger has excellent performance, and the experimental results are in good agreement with the design values, which verifies the feasibility of the new design method.
... In this paper, the inert mass fraction is set to 0.15 for carrying stage and 0.1 for upper stage, which are based on data collected in the literature review. 19,36 Payload mass fraction is defined as payload mass to gross takeoff mass ratio, which is another important parameter to assess the performance of launch vehicle. ...
Article
The liquid oxygen/methane staged cycle liquid-rocket engine is one of the most potential rocket engines in the future for its higher performance, higher fuel density and reusable capacity. Two working states of this liquid-rocket engine named as full-load state and half-load state are defined in this paper. Based on this liquid-rocket engine, a dual-rocket-based-combined-cycle propulsion system with liquid oxygen /air/methane as propellants is therefore proposed. The dual-rocket-based-combined-cycle system has then five working modes: the hybrid mode, pure ejector mode, ramjet mode, scramjet mode and pure rocket mode. In hybrid mode, the booster and ejector rockets driven by the full-load liquid-rocket engine work together with the purpose of reducing thrust demand on ejector rocket. In scramjet mode, the fuel-rich burned hot gas generated by the half-load liquid-rocket engine is used as fuel, which is helpful to reduce the technical difficulty of scramjet in hypersonic speed. The five working modes of dual-rocket-based-combined-cycle are highly integrated based on the full- or half-load state of the liquid oxygen/methane staged cycle liquid-rocket engine, and the unified single type fuel of liquid methane is adopted for the whole modes. Then a preliminary design of a horizontal takeoff two-stage-to-orbit launch vehicle is conducted based on the dual-rocket-based-combined-cycle propulsion system. Under an averaged baseline thrust and specific impulse, the launch trajectory to reach a low Earth orbit at 100 km is optimized via the pseudo-spectral method subject to maximizing the payload mass. It is shown that the two-stage-to-orbit vehicle based on the dual-rocket-based-combined-cycle can achieve the payload mass fraction of 0.0469 and 0.0576 for polar mission and equatorial mission, respectively. Conclusively, insights gained in this paper can be usefully applied to a more detailed design of the dual-rocket-based-combined-cycle powered two-stage-to-orbit launch vehicle.
... Nowadays, special media turbomachinery are increasingly used in various engineering fields. For instance, the helium turbomachinery can be applied to the modular high-temperature gas-cooled reactors (MHTGR) power generation system [1][2][3][4], propulsion system for precooled hypersonic [5][6][7] and superconducting system [8][9]. The supercritical carbon dioxide (SCO2) turbomachinery can be applied to the electric power generation system [10][11]. ...
Article
In order to obtain performance characteristics of special media turbomachinery conveniently and accurately, it is essential to study the similarity method for turbomachinery with different working media. Based on kinematic similarity, a new similarity method, consisting of a new similarity criterion and performance transfer methods, is presented in this paper. In order to validate the method, helium turbomachinery are investigated and air is used as the substitute working medium to simulate the helium turbomachinery. Both the velocity triangle and flow field are contrasted between helium and air turbomachinery when they achieve the similarity criterion derived. Performance map of helium turbomachinery obtained by the performance transfer methods is also compared with that obtained by numerical simulation. Results show that, the new similarity method proposed can be applied to obtaining performance characteristics of special media turbomachinery accurately.
... Inlet air cooling can offer a great technical potential to expand the flight envelope of the turbine engine. Three representative pre-cooling schemes include fuel pre-cooling such as Liquid Air Cycle Engines (LACE) [11] and Deeply Cooled Air Turborocket (ATRDC) Engine [12,13], third-fluid cooling such as Synergetic Air-Breathing Rocket Engine (SABRE) [14], and Mass Injection and Precompressor Cooling (MIPCC) [15]. In allusion to the MIPCC scheme, however, the principle of MIPCC is to evaporative cooling the inlet airflow by injecting a fluid in front of the compressor. ...
Article
The precooled turbine engine is applied to overcome the limitation of Mach number due to high temperature inlet air. This paper aims to investigate the effect of water injection cooling on the high-temperature intake air. Then, the theory evaluation and Eulerian-Lagrangian multiphase flow method are conducted to explore the thermodynamic process and resistance characteristics of the pre-cooling section built-in even spray apparatus with a drag reduction. Results show that larger amount of the injection flow rate at higher Mach number will deteriorate total pressure loss and flow field uniformity. Evaporation cooling can decrease flow loss by 9.4%–60.7%. Within 27 ms, total-temperature drop is in 14–144 K range with a low total-pressure drop coefficient of 0.56%–1.29%. Especially, mass flow will increase by 1.15%–18.50%. Thus, water injection cooling is conducive to a higher acceleration, as well as for improving the thermodynamics characteristics of inlet air for a turbine engine at a high Mach number.
... More recently, Reaction Engines LLC has developed a combined cycle engine that avoids many of the problems with other concepts by cooling the incoming air without condensing it before feeding it into a conventional rocket engine. There has been a great deal of R&D done on this concept [2], [3]. However, this engine requires liquid hydrogen fuel, which is very low in density and requires large tanks to store. ...
... Synergies with composite piston topping could also provide benefits. As a more radical approach, integration with heat transfer systems using a secondary fluid system [19], as well as integration with inter-turbine reheat may be considered. ...
Conference Paper
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Commercial transport fuel efficiency has improved dramatically since the early 1950s. In the coming decades the ubiquitous turbofan powered tube and wing aircraft configuration will be challenged by diminishing returns on investment with regards to fuel efficiency. From the engine perspective two routes to radically improved fuel efficiency are being explored; ultra-efficient low pressure systems and ultra-efficient core concepts. The first route is characterized by the development of geared and open rotor engine architectures but also configurations where potential synergies between engine and aircraft installations are exploited. For the second route, disruptive technologies such as intercooling, intercooling and recuperation, constant volume combustion as well as novel high temperature materials for ultra-high pressure ratio engines are being considered. This paper describes a recently launched European research effort to explore and develop synergistic combinations of radical technologies to TRL 2. The combinations are integrated into optimized engine concepts promising to deliver ultra-low emission engines. The paper discusses a structured technique to combine disruptive technologies and proposes a simple means to quantitatively screen engine concepts at an early stage of analysis. An evaluation platform for multidisciplinary optimization and scenario evaluation of radical engine concepts is outlined.
... The hypersonic precooled engine [1][2][3], which is regarded as one of the potential propulsion methods for the reusable HO-TOL (Horizontal Takeoff and Landing) hypersonic vehicle [4][5][6][7], has attracted great attention in the research field of hypersonic propulsion. In contrast to the conventional aero turbine engine, helium is used in the hypersonic precooled engine to cool down the stagnated air that is of high temperature, forcing the airflow temperature through the inlet of the compressor go back to the level at which the materials and lubricating system can endure [8][9][10]. ...
... 29 This approach was used to design the HOTOL propulsion system−the Rolls Royce RB545 engine−which was a combined cycle pre-cooled engine. 30 The conceptual, unpiloted, reusable Skylon aerospace plane−an SSTO RLV, which can be used as a reusable first stage of a TSTO launch system 31 −(1989-present) evolved from HOTOL. For this concept, various propulsion concepts-hydrogen/oxygen rockets, scramjets, turbojets, turborockets, and Liquid Air Cycle Engines-were considered. ...
... Since the C1 design was frozen REL concentrated on designing, manufacturing and testing critical propulsion system components in wind tunnels and engine test stands. These activities were necessary in order to reduce the risk associated with critical propulsion system components requiring new and unique technologies [8]. This work has been carried out using private investment and recently has received additional funding from ESA and the UK Government. ...
Article
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SKYLON is a reusable single stage to orbit spaceplane intended to lower the cost of reaching space. The project has a 25 year history stretching back to the British Aerospace HOTOL study and over the many configuration iterations the performance has been established using feasibility designs, with market studies being used to establish that the resulting system has utility. In preparation for the final concept study of the D1 configuration the market and other stakeholder's requirements have been prepared as an input rather than an output to the design process. These requirements have been established from both an analysis of the existing market - as a model for the entry into service requirements - and future studies of advanced applications - as a model for the longer term requirements. The final conclusions have been incorporated into a preliminary User Manual which is the basis of a requirements' validation exercise.
... Compared with the above schemes, the precooling combined cycle engine termed SABRE proposed by REL in the United Kingdom represents a substantial technical innovation by introducing helium as the third fluid into the thermal cycle between the incoming airstream and the liquid hydrogen. SABRE has attracted widespread attention and a series of studies have been conducted 4,7,8,[26][27][28] . Advanced compact heat exchangers play an important role in SABRE mainly as a precooler between the air and the medium, as a heat exchanger between the high-temperature gas and the heat-exchange medium, and as a hydrogen/helium heat exchanger. ...
Article
Full-text available
The Hypersonic Precooled Combined Cycle Engine (HPCCE), which introduces precooler into traditional hypersonic engine, is regarded as the most promising propulsion system for realizing a single-stage-to-orbit vehicle. The unique demands lead to the application of the compact heat exchangers, which can realize high thrust-to-weight ratio, sufficient specific impulse and high compression ratio. However, it is challenging to accurately manufacture the compact heat exchanger due to its extremely high heat dissipation capacity, remarkable compactness, superior adaptability and harsh operating condition. This review summarizes the precooling schemes of combined cycle propulsions and describes the demands and key issues in the fabrication of a compact heat exchanger for HPCCE. The investigation focuses on the application of various micromanufacturing methods of heat exchangers constructed from tubes of less than 1 mm in diameter and microchannels of less than 200 micrometers. Various micromanufacturing processes, which include microforming, micromachining, stereolithography, chemical etching, 3D printing, joining and other advanced microfabricating processes, were reviewed. In addition, the technologies are compared in terms of dimensional tolerance, material compatibility, and process applicability. Furthermore, the boundaries of the micromanufacturing constraints are specified as references for the design of compact heat exchangers. Ultimately, the technological difficulties and development trends are discussed for the fabrication of compact heat exchangers for HPCCE.
... National space programs from several countries such as members of the European Union have successfully developed TPS for hypersonic vehicles. This include Germany through its Saenger vehicle [121] and UK through its Skylon vehicle [122]. Other countries such as India, France, Australia, and Japan have all proposed the development of hypersonic capabilities and as such will require TPS for the vehicles [123]. ...
Article
Thermal Protection System (TPS) is an essential component of space vehicles to protect them against the aerothermal heating imposed on them during the atmospheric entry. Different types of TPS, including passive, semi-passive and active systems, have been developed and utilized in the last few decades. With the increasing demand for reusable launch vehicles (RLV) as well as new goals for interplanetary manned missions, the quest for developing effective TPS has been accelerated. This paper provides a comprehensive survey on the technology development of different classes of TPS from mid twentieth century to the present time. Application of different types of TPS for various RLVs is reviewed and the current state of the art of TPS technology is presented. Three major aspects, including mass efficient TPS materials and technology, modeling and simulation tools and techniques and TPS sensors and measurement systems are identified as current challenges pertaining TPS for future space missions, according to the most recent NASA Technology Roadmap. A detailed discussion on these challenges as well as insights regarding future prospects for different classes of TPS are presented in detail.
... The circulating system of SABRE is heat exchanged with the incoming air by introducing helium and liquid hydrogen. Although the engine weight is increased by introducing the helium cycle, the performance is significantly improved in the air-breathing mode owing to the high heliumspecific heat ratio and reduced cycle pressure ratio [47]. ...
Article
Full-text available
The precooled combined cycle engines were proposed to overcome the limitation of Mach number due to high-temperature inlet. However, there has been little discussion about the thermodynamic cycle of these engines. Therefore, the current research progress and key technologies in the precooled engine thermodynamic cycle are analyzed and summarized in detail in this study. The main precooled engines, specifically ATREX and SABRE, are compared. The engine precooling cycle can be divided into direct and indirect precooling cycles. By comparing the performance of the direct precooling cycle ATREX engine in the non-precooling and precooling modes, the advantages of this engine in terms of specific thrust, compressor boost ratio, and flight Mach number are examined. The SABRE engine with an indirect precooling cycle not only solves the problem of hydrogen embrittlement of the precooler, but also achieves higher thrust and specific impulse than the ATREX engine. However, the ATREX engine with a direct precooling cycle is simpler in structure and has less adjusting variables, and thus, it is suitable for application in high-speed aircraft with a low flow rate. In addition, comparing the precooling cycle working mediums, the hydrogen medium shows excellent performance but has a limited aircraft flight range. Conversely, the use of a hydrocarbon fuel in the aircraft results in a longer flight range but with relatively low performance. Therefore, a development trend toward multi-fuel hybrid use can be expected.
... Nowadays, the hydrogen powered hypersonic precooled aeroengine, which is one of the most prospective propulsion systems, has attracted extensive concerns in the field of hypersonic transportation [1][2][3]. The feasibility of the whole system depends on the light-weight and high-efficient heat exchanger. ...
Article
The microchannel heat exchanger is a critical component of hypersonic precooled aero-engines, suffering huge challenges from design and manufacture due to the demanding requirements of higher power-weight ratio and lower pressure loss. In this paper, a segmental method (SEG) taking the influence of fin efficiency into consideration is proposed to analyze and evaluate the performance of microchannel heat exchanger for precooled engines. In order to validate the method and investigate the flow and heat transfer characteristic, a microchannel heat exchanger with the hydraulic diameter of 263.6 μm was fabricated and measured on a 3 kW open loop experimental platform. The experimental results indicate that the maximum deviation between the heat transfer rate predicted by the SEG method and the experimental data does not exceed 5% on the condition of balanced flow and unbalanced flow. Furthermore, a new correlation for pressure drops based on experimental results without heat transfer is developed to correct the results predicted by the SEG method and the maximum deviation does not exceed 3%. At last, the influence of the geometric parameter of the basic heat transfer unit was thoroughly discussed for further understand the performance of microchannel heat exchanger under typical working conditions. It is indicated that potential channel width should not be more than 0.5 mm (= 0.305 mm) h d in the case of integrating various factors.
... Launch costs make up a significant fraction of total costs, and could be reduced in the near term by developments, such as SpaceX's reusable first stage [SpaceX, 2013]. However, a SSTO system, such as the proposed Skylon, could drastically reduce launch costs [Bond, and Varvilland, 2008]. Utilizing reusable launch infrastructure also has the potential to reduce the environmental impact of AMOOS. ...
Article
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This paper presents the conclusions of the Team Project "Autonomous Mission for On-Orbit Servicing" (AMOOS) of the Space Studies Program 2014 (SSP'14) from the International Space University (ISU), held at the campus of the École de Technologie Supérieure (ETS) in Montreal, Quebec, Canada. The AMOOS project aims to conduct a multidisciplinary feasibility analysis of robotics-equipped, autonomous space planes to service satellites and to remove debris in Low Earth Orbit (LEO). The five main objectives of the Team Project are: (1) to identify and formulate future key space technologies of unmanned spacecraft and robotic systems suitable for an autonomous on-orbit servicing, (2) to understand the space debris issue and to examine the feasibility of orbital debris removal with an unmanned space plane and embedded robotic systems, (3) execute orbital maneuvers simulations and analyze the results, (4) study the economic, legal, environmental, and social aspects of Active Debris Removal (ADR) and On-Orbit Servicing (OOS) missions, and (5) to engage the public, space agencies, space industries and government officials about the space debris issue. The continuity of current space operations raises two essential points of discussion: the extension of operational life of a satellite in orbit and the reduction or removal of space debris. To address these issues, the Team Project analyzed, developed, and discussed the utilization of space planes and robotic systems to: (a) service satellites to extend their lifetime and performance, (b) deploy small and secondary payloads in Low Earth Orbit, and (c) capture and/or de-orbit large space debris and non-operational satellites. To achieve the listed objectives, the Team Project investigated available technologies and assessed the technical feasibility of AMOOS by demonstrating the potential economic benefits. Finally, to demonstrate political and legal benefits, the team Project investigated potential threats and risks related to ADR and OOS. Copyright © 2014 by the International Astronautical Federation. All rights reserved.
Article
A review of the key technologies of air-breathing engine precooling system and the use of micro-channel heat transfer is presented. A survey on various types of air-breathing engine cycles is presented, highlighting the characteristics of the energy cycles and the corresponding key technologies. The existing precooling schemes are classified into four types, i.e. Fuel Precooling (FPC), Mass Injection and Precompressor Cooling (MIPCC), combination of FPC and MIPCC, and Third-Fluid Cooling (TFC). Precoolers with micro-channel structures are found to have high heat dissipation capacity and high compactness. In detail, the applications of microchannel flow heat transfer like in the SABRE engine of the British Skylon spaceplane were introduced. Fundamental investigations on the microchannel heat transfer enhancement are essential for the development of the precooling technique. In order to better understand the micro-channel heat transfer mechanisms, experimental studies on single phase gaseous flow heat transfer in small flow passages are briefly overviewed, revealing some controversial conclusions on microscale flow and heat transfer characteristics. The limited experimental data on microchannel gaseous flow heat transfer largely hinders the theoretical development. Since the air precooling technique is at its infancy in China, experimental investigations are essential to overcome the gap.
Article
Mass injection pre-compressor cooling (MIPCC) is one of the significant technologies to extend the operating envelope of turbine engines. In this paper, a new high Mach number turbine scheme is proposed with liquid ammonia injection instead of water injection. The basic properties of ammonia, especially the boiling point, latent heat of vaporization and auto-ignition temperature, are introduced. To evaluate the engine performance, a thermodynamic model incorporating the ammonia mass injection pre-compressor cooling (Ammonia MIPCC) process is developed. The accuracy of the model is demonstrated by verifying the design point, operating line, and Ammonia MIPCC processes. Simulations show that the Ammonia MIPCC process is mainly suitable for extended operating Mach number envelopes. Ammonia injection should be avoided for normal operating conditions because it causes a decrease in specific impulse. Through liquid ammonia injection, the operating Mach number of the turbine engine is increased from 3.1 to about 3.5. Moreover, the maximum operating Mach number (MOM) limit of the Ammonia MIPCC engine is investigated. Pre-Compressor cooling requirements and combustor heat release limits jointly constrain the amount of ammonia injection, which prevents further expansion of the MOM. Furthermore, the Ammonia MIPCC engine has the advantage of greater specific thrust at the same Mach number compared to water injection, and with Mach numbers greater than 3.36, the Ammonia MIPCC engine has a specific impulse advantage. In general, the Ammonia MIPCC engine is an effective solution for a high Mach number turbine.
Conference Paper
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Re-usable single stage to orbit launch vehicles promise to reduce the cost of access to space, but their success will be particularly reliant on accurate and robust modelling of their aero-thermodynamic characteristics. For preliminary design and optimization studies, relatively simple numerical prediction techniques must perforce be used, but it is important that the uncertainty that is inherent in the predictions of these models be understood. Predictions of surface pressure and heat transfer obtained using a new reduced-order model that is based on the Newtonian flow assumption and the Reynolds analogy for heating are compared against those of a more physically-sophisticated Direct Simulation Monte Carlo method in order to determine the ability of the model to capture the aero-thermodynamics of vehicles with very complex configuration even when run at low enough resolution to be practical in the context of design optimization studies. Attention is focused on the high-altitude regime where lifting re-usable Single-Stage to Orbit configurations will experience their greatest thermal load during re-entry, but where non-continuum effects within the gas of the atmosphere might be important. It is shown that the reduced-order model is capable of reproducing the results of the more complex Monte Carlo formalism with surprising fidelity, but that residual uncertainties exist, particularly in the behaviour of the heating models and in the applicability of the continuum assumption given the onset of finite slip velocity on surface of vehicle. The results suggest thus that, if used with care, reduced-order models such as those described here can be used very effectively in the design and optimization of space-access vehicles with very complex configuration, as long as their predictions are adequately supported by the use of more sophisticated computational techniques.
Article
The presence of asymmetric side loads due to unstable separation within over-expanded rocket nozzles is well documented. Although progress has been made in developing understanding of this phenomenon through numerical and experimental means, the causes of these side loads have yet to be fully explained. The hypothesis examined within this paper is that there is a relationship between nozzle wall angle at the point of separation, and the stability of the flow separation. This was achieved through an experimental investigation of a series of subscale over-expanded conical nozzles with half-angles of 8.3°, 10.4°, 12.6° and 14.8°. All had overall area ratios of 16:1, with separation occurring at approximately half the nozzle length (i.e. area ration of 4:1) under an overall pressure ratio of approximately 7:1 using air as the working fluid. The structure of exhaust flow was observed and analysed by use of an optimised Schlieren visualisation system, coupled with a high speed digital camera. The 12.6° and 14.8° nozzles exhaust flow were seen to be stable throughout the recorded test period of 10 seconds. However, a small number of large fluctuations in the jet angle were seen to be present within the flowfield of the 10.4° nozzle, occurring at apparently random intervals through the test period. The flowfield of the 8.3° nozzle demonstrated near continuous, large angle deviations in the jet, with flow patterns containing thickened shear layers and apparent reattachment to the wall, something not previously identified in conical nozzles. These results were used to design a truncated ideal contour with an exit angle of over 10 degrees, in order to assess the possibility of designing conventional nozzles that separate stably over a wide range of pressure ratios. These tests were successful, potentially providing a simpler, cheaper alternative to altitude compensating nozzle devices. However, more work determining the nature of the separation and its causes is required.
Article
The SABRE engine for SKYLON has a sophisticated thermodynamic cycle with heat transfer between the fluid streams. The intake airflow is cooled in an efficient counterflow precooler, consisting of many thousand small bore thin wall tubes. Precooler manufacturing technology has been under investigation at REL for a number of years with the result that flightweight matrix modules can now be produced.A major difficulty with cooling the airflow to sub-zero temperatures at low altitude is the problem of frost formation. Frost control technology has been developed which enables steady state operation.The helium loop requires a top cycle heat exchanger (HX3) to deliver a constant inlet temperature to the main turbine. This is constructed in silicon carbide and the feasibility of manufacturing various matrix geometries has been investigated along with suitable joining techniques.A demonstration precooler will be made to run in front of a Viper jet engine at REL's B9 test facility in 2011. This precooler will incorporate full frost control and be built from full size SABRE engine modules. The facility will incorporate a high pressure helium loop that rejects the absorbed heat to a bath of liquid nitrogen.
Conference Paper
A preliminary study to assess the existence of shock-shock interactions which might jeopardize the SKYLON structure indicated that the junction of nacelle and wing is the most critical region of the vehicle. Therefore, within the present paper the potential of effusion cooling with steam to control the heat flux in these regions of SKYLON is investigated. The simulations with the three-dimensional finite volume Euler/Navier-Stokes solver TAU investigate the most critical non-equilibrium hypersonic flow conditions along the SKYLON re-entry trajectory. They underline the feasibility of effusion cooling to protect critical structure regions of spacecraft.
Article
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SKYLON is a reusable single stage to orbit spaceplane that can take off from a runway, reach a 300 km altitude low earth orbit with a payload of 15 tonnes and then return to earth for a runway landing. The feature that enables this is the Synergistic Air- Breathing Rocket Engine (SABRE) which has both air breathing and pure rocket modes. The project has been conceived as a commercial venture with the objective that the price charged for the launch, covers all operational and acquisition cost with profit. That means access to space becomes a pure economic activity without the need for public subsidy of the development or day to day running costs of the launch activity. A key way to achieve this objective is the separation of the supplier of the SKYLON system and the operator, following the model in the air transport industry where airliner manufacturers build aircraft that are then sold to many different competing airlines. This approach allows commercial development operations without any assumptions about growth in the market for space launches.
Conference Paper
This paper presents preliminary results for the local heat transfer distributions around a circular cylinder buried deep within a compact heat exchanger matrix. This forms the initial phase of a research program, sponsored by Reaction Engines, to enhance the heat transfer performance of their precooler compact heat exchanger matrix. This high performance matrix forms a critical component of the unique hybrid airbreathing and rocket engine cycle, known as SABRE, that is being developed in order to reduce the cost of access to space by enabling the realization of a fully reusable single stage to orbit spaceplane (Skylon). The performance enhancement of this heat exchanger (whose role it is to provide precooling of the airflow into the SABRE engine) via conventional methods i.e. compactness, wall thickness, flow arrangement etc has reached its practical limit. The substantial gains in payload offered, however, by further performance enhancement using alternative methods provides the motivation for this current research program. © 2010 by Reaction Engines Limited. Published by the American Institute of Aeronautics and Astronautics, Inc.
Conference Paper
The programming environment EcosimPro together with the European Space Propulsion Simulation System has been used to model a high speed propulsion system. The selected engine is the Synergistic Air- Breathing Rocket Engine (SABRE), which is the power plant of the Skylon space plane. The particularities of this engine comprise the use of a pre-cooler in the accelerating air-breathing phase. By contrast, in the anaerobic phase the engine behaves as a staged combustion rocket. New models for the turbomachinery components based in experimental performance maps have been developed. This ad hoc compressor, turbine and circulator components are evaluated at off-design conditions. A full simulation has been done to compute the SABRE performance during its air-breathing trajectory from take-off to Mach 5 at 25 km altitude. This analysis yields the variation of specific impulse and thrust necessary to design high speed vehicles. This model is being coupled by an optimizer to allow the definition of the optimal operating conditions for a given trajectory. Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc.
Conference Paper
The paper will describe the system concept of the SKYLON Single Stage to Orbit Reusable Launch Vehicle. The basic mission will be described, and a technical description of the propulsion system and the airframe will be presented. Current activities, and the experimental test programme that is verifying the critical technologies, will be discussed. In particular, the design, manufacture and test of the only item requiring brand new technology - the precooler - will be discussed, and the consequent risk reduction activities placed into context. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
Chapter
Launchers and reentry vehicles have in common the necessity to perform part of their trajectory along the planet atmosphere. While this has only negative effects for the launch vehicle, for the re-entry vehicle, the interaction with the atmosphere can be suitably exploited for the fulfillment of the mission. The launch and reentry trajectory is a complex scenario that should be modeled using simplified physics equations describing with sufficient accuracy the subsystems of the vehicle and the environment. In a second step, this simplified physical model, described by sets of differential equations, should be transcribed in a mathematical set of algebraic equations that can be solved by non-linear programming methods (NLP solvers). Upon further analysis of these direct transcription methods, two subclasses can be identified: shooting methods and collocation methods. The NLP solver can be a global optimizer, e.g., genetic algorithm, particle swarm, some other metaheuristics or sequential quadratic programming (SQP), in the case of differentiable functions. Several examples of launch and reentry vehicle problems are presented, with a strong emphasis on the advantages and disadvantages of the various transcription methods and solvers when applied to “real” world problems.
Article
Comparing advanced reusable spaceplane concepts and programmes, Buffo identified seven attributes to describe a true spaceplane. This paper adds two more attributes to clearly distinguish spaceplanes from rocketplanes. One of them is the hydrogen fuel fraction (HFF), the ratio of the mass of liquid hydrogen fuel to the take-off mass. Excellent correlation between HFF and spaceplane techno-economic performance is seen from regression trends arising from parametric analysis. A modified rocket equation, termed spaceplane equation, shows how high HFF, scalability and multi-stage rocket performance are feasible in single stage to orbit (SSTO) spaceplane by the combined use of fuel efficient airbreathing engines and in-flight LOX addition (FLOX). A LACE-FLOX aerocryogenic engine cycle with air liquefaction (using LACE engines) integrated with vortex/higee liquid oxygen separators may well provide both high HFF and a single combined-tycle engine from earthto-orbit. The global status of aerocryogenic technologies is also provided.
Article
The term combustible inertial with reference to a launch vehicle means that its case is used as main propellant during a flight being fed into a rocket engine by means of inertial forces only (the engine has no turbopump or gas-pressure feed system). Theoretically this technology allows developing extremely light-weight (several hundred kilograms of initial mass) quasi single-stage solid-propellant launch vehicle capable of putting 1 kg payload into a low Earth orbit. The simple structure of the launch vehicle enables assembling, servicing and launching of it by a small group of individuals. The version of the launch vehicle design is shown as dimensioned pictures. Results of calculations and previous experiments are presented. The next step, the experimental investigation of a laboratory propulsion transferring into launching it as a small rocket, is quite accessible for a small university or amateur team. Being successful, this can answer in the affirmative the question of the paper title.
Article
This study explores the design, analysis, and air pressure drop assessment of three analogous air-fuel heat exchangers consisting of thin serpentine tube bundles intended for use in high Mach number aero-engines. In high speed flight, the compressor bleed air used to cool high temperature turbine blades and other hot components is too hot. Hence, aviation kerosene is applied to precool the compressor bleed air by means of novel air-fuel heat exchangers. Three light and compact heat exchangers including dozens of in-line thin serpentine tube bundles were designed and manufactured, with little difference existing in aspects of tube pitches and outer diameters among three heat exchangers. The fuel flows inside a series of parallel stainless serpentine tubes (outer diameter: 2.2, 1.8, 1.4 mm with 0.2 mm thickness), while the air externally flows normal to tube bundles and countercurrent with fuel. Experimental studies were carried out to investigate the airside pressure drop characteristics on isothermal states with the variation of air mass flow rates and inlet temperatures. Non-isothermal measurements have also been performed to research the effect of heat transfer on pressure drops. The experimental results show that inlet temperatures have significant influence on pressure drops, and higher temperatures lead to higher pressure drops at the same mass flow rate. The hydraulic resistance coefficient decreases quickly with Reynolds number, and the descent rate slows down when Re>6000 for all three heat exchangers. Additionally, the pressure drop on heat transfer states is less than that on isothermal states for the same average temperatures. Moreover, the pressure drop through heat exchangers is greatly affected by attack angles and transverse pitches, and an asymmetric M-shaped velocity profile is generated in the cross-section of sector channels.
Article
The Printed Circuit Heat Exchanger (PCHE) is one of the most promising heat exchangers for Synergetic Air-breathing and Rocket Engine (SABRE). To reduce pressure drop and improve compactness, the micron-sized PCHE made up of rectangular channels of tens of microns in size, is used in SABRE. In present work, we focus on thermal-hydraulic-structural characteristics of micron-sized PCHE by conducting three-dimensional (3-D) numerical simulation. Helium and hydrogen are employed as the working fluids and the Stainless Steel 316 (SS316) as the solid substrate. The thermal-hydraulic performance of the micron-sized PCHE is discussed by using the commercial Computational Fluid Dynamics (CFD) software of Fluent. ANSYS-Mechanical is also employed to simulate stress field of representative PCHE channels. The mechanical stress induced by pressure loading and the thermal stress induced by temperature gradient are found to be equally important sources of stress. To improve comprehensive performances of micron-sized PCHE, two types of channel arrangements and different channel aspect ratios are studied. The double banking is of higher thermal-hydraulic performance compared to the single banking while the stress performance is identical for the two modes. Meanwhile, the effect of channel aspect ratio is investigated by comparing thermal-hydraulic characteristics and structural stress of the model. The rectangular channel with w/h=2 achieves the most balanced stress characteristic and higher thermal-hydraulic performance.
Article
Fuel indirect precooled engines have the potential to radically reform the paradigm of propulsion for next generation in- and trans-atmospheric vehicles. The engine is characterized by a sophisticated thermodynamic cycle with a series of heat exchangers. Reducing irreversibility of the precooling-compression system (PCS) is the key to ensure high engine performance, for which a conjugated optimization coupled with heat exchanger design was performed to evaluate the performance level that can be achieved under the state of technologies. The results indicate that irreversibility of the PCS can be reduced but at the penalty of heavier and larger size of heat exchangers. Moreover, it shows that precooler and regenerator together contributes 75–84% of the total entropy generation rate within the PCS, while hydrogen pump occupies the second largest irreversibility source. Smaller precooling temperature can help to improve the extent of irreversibility of the PCS and increase the maximum achievable air pressure ratio and engine specific thrust, nevertheless the increased fuel consumption and air side pressure drop and the decreased engine specific impulse shouldn’t be ignored. Larger regenerator effectiveness is preferred from all aspects of the PCS and engine level figures of merit as long as the total weight and length of the heat exchangers are not increased aggressively.
Article
Precooled engine is a highly expected solution to achieve supersonic transport. As the crucial component, heat exchangers protect other components from ultrahigh temperature. In traditional design methods, the nominal result is multiplied by a safety factor, whose selection entirely depends on experience, ensuring sufficient working margin to cope with fluctuation of parameters. For aero-engine, heat exchangers must work reliably with minimum weight. An advanced method of thermal optimization design with parameters’ fluctuation is proposed and proved to be effective by experimental verification. The heat transfer area can be quantitatively linked with the design confidence level, considering the coupling effect of various parameters’ fluctuation. The probability density distribution of heat transfer area has the characteristic of positive skewness distribution. With the increase of design confidence, the required heat transfer area is growing faster and faster. After optimization, the design of heat exchanger meets the requirements and the weight is effectively controlled.
Article
The rapid change in the amount of carbon dioxide released into the atmosphere stimulates the advancement of advanced space launchers. Liquefied air cycle engine is a pollution-free space launcher and attracts much attention in recent years. Air-hydrogen precooler is an important component in the liquefied air cycle engines. The thermal–hydraulic performance of the precooler significantly affects the efficiency of the engine. This paper proposes a novel annular air-hydrogen precooler based on printed circuit heat exchanger. To prove the thermal–hydraulic performance of the novel precooler design, three types of precoolers are selected for comparison of volumetric power, power per mass unit and compactness. Segmented thermal design method and genetic algorithm are used for structure optimization. The results indicate that the novel precooler has higher compactness and volumetric power. The compactness of annular air-hydrogen precooler is 682% larger than that of shell-and-tube heat exchanger. The volumetric power of annular air-hydrogen precooler is 206% larger than that of plate-fin heat exchanger. The volumetric power of annular air-hydrogen precooler is 24% larger than that of the printed circuit heat exchanger. The study concludes that the air-hydrogen precooler has a better comprehensive performance.
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The fuel indirect precooled turbine combined cycle engine is the promising power to realise more green environmental protection intercontinental navigation and hypersonic aerospace transportation in the future. The compact, lightweight precooler arranged in the front of the inlet is a critical component of the advanced heat management system, which could improve the fuel efficiency and flight range of the turbine engine. This paper mainly focuses on the high-temperature brazing process of Inconel 718 ultrathin-walled capillary-and-plate brazed structure using the BNi-5 brazing filler metal. In this investigation, a testing method for tensile strength of ultrathin-walled structures was proposed; the effects of different brazing parameters on the mechanical properties and microstructure evolution of different types of brazed structures were studied to investigate the particularity of ultrathin-walled capillary-and-plate brazed structure. The brazed joint was mainly composed of γ-Ni solid solution, Ni5Si2, G-phase, Ni3Si, and Cr3Ni5Si2. The results showed that the higher brazing temperature and longer holding time were beneficial to improve the mechanical properties of the conventional lap brazed structure; however, the mechanical properties of the ultrathin-walled capillary-and-plate structure decrease significantly. The optimised brazing process for the ultrathin-walled structure was 1150 °C for 3 min; dissolution behaviour of the ultrathin-walled capillary and the deeper position of the precipitates induced by the acute diffusion phenomenon play an important role in the tensile strength of the ultrathin-walled structure. And the failure mechanism was analysed in detail; the result indicated that ultrathin-walled capillary-and-plate brazed structure was mainly affected by the eutectic structure of the brazing fillet; the Nb-rich precipitates in the diffusion affected zone and the dissolution behaviour of the ultrathin-walled base metal.
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Porous Ultra-High-Temperature-Ceramics (UHTC) are a candidate group of materials for transpiration cooling of hypersonic vehicles due to their exceptionally high melting point, typically above 3000 K. Their high operating temperature permits a higher amount of radiative cooling than that achievable with conventional materials, which reduces the required coolant mass flow rate to cool the surface. This work experimentally examines the internal and external flow behaviour of porous UHTC made of zirconium diboride (ZrB2) for the purpose of transpiration cooling. A dedicated ISO standard permeability test rig was built. The outflow velocity distribution was acquired employing miniature hot-wire anemometry. The data obtained for the pressure loss across the porous samples agree with the Darcy-Forchheimer model for flow in porous media; respective Darcy and Forchheimer permeability coefficients are calculated and reported. Cleaning the surface of the samples using sandpaper or an ultrasonic bath raised the permeability coefficient by up to 19%. The outflow velocity maps exhibit a good flow uniformity with an average standard deviation of 25.1% with respect to the mean value. Individual jets are absent, and the velocity varies within the same order of magnitude.
Chapter
Thermal protection systems (TPSs) for space and hypersonic vehicles encompass a wide range of materials and design designs. In any particular vehicle, a variety of materials may be used. The choice of materials and thermal protection concepts is based upon thermal environment and exposure times, as well as whether the application requires single use or multiuse. This chapter provides an overview of materials and concepts, including tile, blankets, metallic systems, structurally integrated concepts, ceramic matrix composites (CMCs), used primarily in hot structures, ultra-high temperature ceremics (UHTCs), and ablators. It is based upon a review of publicly available literature and includes a historical perspective, as well as some newer developments as of 2009.
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Expansion Deflection nozzles present an attractive proposition as a replacement for conventional nozzles on launch vehicles, due to their reduced length, and altitude compensating capability. However, it has long been speculated that they suffer in the latter regard due to aspiration of the low speed flow region inside the nozzle by the supersonic jet surrounding it. This effect is investigated in this paper by direct experimental measurement of base pressures, and found to have little effect on the base pressure of the nozzle within the range of operating conditions investigated. Wall pressures were also used to calculate the efficiency of the altitude compensation within the nozzle, which was found to be between 87 and 100% for the three operating pressure ratios examined. This represents a significant improvement over conventional nozzle performance, and further conformation that wake pressures are indeed close to ambient.
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This paper considers issues relevant to propulsion design for Single Stage To Orbit (SSTO) vehicles. In particular two air-breathing engine concepts involving precooling are considered, these being the SABRE (Synergetic Air-Breathing and Rocket Engine) as designed for the SKYLON SSTO launch vehicle, and the LACE (Liquid Air Cycle Engine). It is shown that through entropy minimization the SABRE has made substantial gains in performance over the traditional LACE precooled engine concept, and has shown itself as the only credible means of realising a SSTO vehicle. Furthermore, it is demonstrated that the precooler is a major source of thermodynamic irreversibility within the engine cycle and that a further reduction in entropy can be realised by increasing the heat transfer coefficient on the air side of the precooler. If this may be achieved, an increase of between 5 and 10% payload mass delivered to orbit by the SKYLON launch vehicle is possible.
Article
The issues relevant to propulsion design for Single Stage To Orbit (SSTO) vehicles are considered. In particular two air-breathing engine concepts involving precooling are compared; SABRE (Synergetic Air-Breathing and Rocket Engine) as designed for the Skylon SSTO launch vehicle, and a LACE (Liquid Air Cycle Engine) considered in the 1960's by the Americans for an early generation spaceplane. It is shown that through entropy minimisation the SABRE has made substantial gains in performance over the traditional LACE precooled engine concept, and has shown itself as the basis of a viable means of realising a SSTO vehicle. Further, it is demonstrated that the precooler is a major source of thermodynamic irreversibility within the engine cycle and that further reduction in entropy can be realised by increasing the heat transfer coefficient on the ah- side of the precooler. If this were to be achieved, it would improve the payload mass delivered to orbit by the Skylon launch vehicle by between 5 and 10%.
Article
While expansion deflection (ED) nozzles have traditionally been considered primarily for use as altitude compensating devices to improve the performance of single stage to orbit vehicles, they also offer the potential for enhancing high altitude propulsion systems. If intended to only operate in near vacuum conditions, the complexity of analysis and inherent risks involved in the ED concept are greatly reduced. An integrated approach to the design and performance analysis of such nozzles is presented, comprising a mixture of computational fluid dynamics, the method of characteristics, and a semi-empirical model to allow full analysis of the closed wake flow-field of an ED nozzle. While it is demonstrated that the influence of the parameters used to define the throat region is critical to the successful application of the ED nozzle, it is also shown that with careful design the weight savings possible are significant. The analysis method itself is flexible and rapid, and lends itself well to incremental improvements in accuracy as the flow under consideration becomes better understood.
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A small pre-cooling heat exchanger with 0.38 mm tubes has been built and tested using conditions reflecting the operating characteristics of the helium flight precooler in the air breathing rocket engine of the SKYLON aerospaceplane. This paper discusses the manufacturing techniques developed for the construction of fine tube high pressure heat exchangers with compactnesses up to 3000 m2/m3. Work centred on brazing large numbers of very thin wall tubes to high reliability. Repeatability of the results was ensured by the use of a plating technique for pre-placement of the braze alloy. Process conditions were also refined to facilitate good wetting of high chromium steels with very low erosion on thin wall tubes in standard industrial vacuum furnaces. The paper also discusses the results obtained from heat transfer experiments using the exchanger. Power transfer rates close to 2 GW/m 3 with temperature changes of up to 500°C were observed. The results showed that heat transfer performance remained predictable down to hydraulic diameters of a third of a millimetre. The paper also discusses inconsistencies that were encountered in the matrix pressure loss predictions. These could not be explained using the one dimensional correlations currently employed.
Article
This paper discusses the relevant selection criteria for a single stage to orbit (SSTO) propulsion system and then reviews the characteristics of the typical engine types proposed for this role against these criteria. The engine types considered include Hydrogen/Oxygen (H2/O2) rockets, Scramjets, Turbojets, Turborockets and Liquid Air Cycle Engines. In the authors opinion none of the above engines are able to meet all the necessary criteria for an SSTO propulsion system simultaneously. However by selecting appropriate features from each it is possible to synthesise a new class of engines which are specifically optimised for the SSTO role. The resulting engines employ 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 modes. This results in a significant mass saving with installation advantages which by careful design of the cycle thermodynamics enables the full potential of airbreathing to be realised. The SABRE engine which powers the SKYLON launch vehicle is an example of one of these so called `Precooled hybrid airbreathing rocket engines' and the concep- tual reasoning which leads to its main design parameters are described in the paper.
Article
A reusable SSTO spaceplane employing dual mode airbreathing/rocket engines, such as SKYLON, has a voluminous fuselage in order to accommodate the considerable quantities of hydrogen fuel needed for the ascent. The loading intensity which this fuselage has to withstand is relatively low due to the modest in-flight inertial accelerations coupled with the very low density of liquid hydrogen. Also the requirement to accommo- date considerable temperature differentials between the internal cryogenic tankage and the aerodynamically heated outer skin of the vehicle imposes an additional design constraint that results in an optimum fuselage structural concept very different to conventional aircraft or rocket practice. Several different structural con- cepts exist for the primary loadbearing structure. This paper explores the design possibilities of the various types and explains why an independent near ambient temperature CFRP truss structure was selected for the SKYLON vehicle. The construction of such a truss structure, at a scale not witnessed since the days of the airship, poses a number of manufacturing and design difficulties. In particular the construction of the nodes and their attachment to the struts is considered to be a key issue. This paper describes the current design status of the overall truss geometry, strut construction and manufacturing route, and the final method of assembly. The results of a preliminary strut and node test programme are presented which give confidence that the design targets will eventually be met.
Article
Modelling of the supersonic flow within a rocket nozzle of both conventional and expansion deflection (ED) design is well handled by Method of Characteristics based algorithms. This approach provides both a predic- tion of the flowfield, and allows efficient optimisation of nozzle shape with respect to length. However, the Method of Characteristics requires a solution of the transonic flow through the nozzle throat to provide initial conditions, and the accuracy of the description of the transonic flow will clearly affect the overall accuracy of the complete nozzle flow calculation. However, it is relatively simple to show that conventional analytical methods for this process break down when applied to the more complex throat geometry of ED nozzles. This requires the use of a time marching solution method, which allows the analysis of the flow within this region even on such advanced configurations. This paper demonstrates this capability, outlines a general method for ED nozzle throat geometric definition, and examines the effect of various throat parameters on the permissible range of ED contours. It is found that the design of length optimised ED nozzles is highly sensitive to small changes in these parameters, and hence they must be selected with care.
Article
An unmanned launch vehicle design is reviewed which specifies single stage to orbit (SSTO), hybrid air-breathing/rocket propulsion, horizontal takeoff and landing, and autonomous, unmanned operation in satellite launch and recapture. The mass of the HOTOL is 80 percent propellant which makes engine efficiency important, and the horizontal take-off requires optimization of the power unit scale. The structural design includes a carbon-PEEK liquid hydrogen tank and a large plan area and a titanium fuselage with carbon-carbon panels for thermal protection, and take-off is accomplished with a reusable launch vehicle. Aerodynamic considerations include minimizing drag, and optimizing intake capture area while preserving reduced drag. Continued safe operation of the systems is possible after two failures, and the unmanned operation includes a constant EAS trajectory followed by a steeper rocket climb.
The SKYLON Spaceplane
  • R Varvill
  • A Bond
R. Varvill and A. Bond, "The SKYLON Spaceplane", JBIS, 57, pp.22-32, 2004.
Design and manufacturing study for a small, complex component required in large production volumes
  • K D Potter
  • A Towse
  • M R Wisnom
K.D. Potter, A. Towse and M.R. Wisnom, "Design and manufacturing study for a small, complex component required in large production volumes", International Conference of Composite Materials, 14-18 July 1997, Australia, IV, pp.103-112.