Modelling of a Dual-Fuel-Mode Free-Jet Combustion System
Modelling of a Dual-Fuel-Mode Free-Jet Combustion System
Cooper, Maxim;Sam, Ashish Alex;Pesyridis, Apostolos
2019-12-17 00:00:00
aerospace Article Modelling of a Dual-Fuel-Mode Free-Jet Combustion System y y ,y Maxim Cooper , Ashish Alex Sam and Apostolos Pesyridis * College of Engineering and Design, Brunel University London, Uxbridge, London UB8 3PH, UK; 1108868@brunel.ac.uk (M.C.); ashish.sam@brunel.ac.uk (A.A.S.) * Correspondence: a.pesyridis@brunel.ac.uk; Tel.: +44-18-9526-7901 y These authors contributed equally to this work. Received: 21 October 2019; Accepted: 10 December 2019; Published: 17 December 2019 Abstract: The focus of this study is to design a combustion system able to sustain hypersonic flight at Mach 8. A Dual-Mode Free-Jet combustion chamber design, first tested in 2010 by NASA, is being adapted to run on hydrogen fuel instead of ethylene while addressing the excessive thermal heat load. This study is part of the FAME (Flight at Mach Eight) project, with the primary objective to design and analyse the engine configuration for a hypersonic commercial aircraft. This CFD analysis and validation study, the first to replicate this combustion chamber design, provides detailed instructions on the combustion system design. The analysis from this study can be used for future research to successfully reach a sustainable design and operation of a Dual-Mode Free-Jet combustion chamber. The 53% size reduction in the combustion system represents significant progress which encourages future research regarding in the design of combustion systems for hypersonic propulsion systems. Keywords: hypersonic aircraft; CFD; dual-mode combustion 1. Introduction The development of any hypersonic aircraft must take into consideration the requirements of all the dierent components that make-up the aircraft and how their designs aect one another, from the lift generation component, fuel, propulsion system and control surfaces [1]. The invention of safe and reliable jet engines made any part of the world accessible within 24 h. The Concorde and the Tupolev Tu-144 have been the only supersonic transport aircraft to see regular service. The first step to making supersonic flight accessible to civilians was the creation of the Concorde flying at Mach 2.02 in 1970. Flying above subsonic speeds, Mach 0.8 to 0.9, is considered inecient, due to the sudden increase of drag when travelling above the speed of sound [2]. Concorde’s retirement in 2003 was due to a combination of inflated fuel prices and the infamous accident in which 113 people lost their lives. The loss of confidence in the safety of the aircraft was the same reason for the Tupolev Tu-144 retirement [3]. Other aircraft have reached speeds greater than Mach 2 but relied upon dierent engine technologies such has ramjets, scramjets and rockets. As these types of propulsion technologies are fundamentally dierent to the standard turbojet and turbofan engines, many technical challenges will have to be overcome before being their introduced to the civil aircrafts. The extreme harsh flow conditions, within the engine as well as the outside of it, required to achieve high Mach velocities need to be very carefully managed for it to become reality. The engine configuration studied in this report has in mind taking the least amount of volume whilst still producing enough thrust for an aircraft for commercial use like the Concorde. Aerospace 2019, 6, 135; doi:10.3390/aerospace6120135 www.mdpi.com/journal/aerospace Aerospace 2019, 6, 135 2 of 18 1.1. Hypersonic Flows The recent development of hypersonic aircraft and guided missiles going faster has brought new engineering challenges. The accepted convention in aerodynamics that defines hypersonic flow is where the flow speeds are greater than Mach 5. At these speeds, the physical behavior of air is very dierent to what is encountered at subsonic. When an object is flying through at the atmosphere, at supersonic or hypersonic speed, a shock layer forms upstream of it, called a bow shock. This phenomenon must be considered due to the abrupt changes in flow properties across it, such as entropy increase, temperature and pressure. The viscosity of the fluid plays an intricate role with the internal energy of the fluid. The high temperatures associated with such flows may generate the dissociation and ionization of air molecules [4,5]. 1.2. Reaching Mach 2.0 The Concorde was powered with four Olympus 593 Mrk610 turbojet engines manufactured by Rolls-Royce, UK and Snecma of France. The thrust produced by each engine at take-o (with afterburner) and during supersonic cruise were 170 kN and 45 kN, respectively. The engines were run on A1 Jet fuel, which is still widely used today. The power plant was fitted with a two-spool system which would run at high and low compression stage, each made of 7 stages and driven by two turbine systems. At cruise speeds, the compression ratio would reach 80:1. The high temperatures resulting from the high compression ratio created a thermal challenge for engineers. Parts of the engine required materials with high service temperatures that only nickel-based alloys could provide at the time. In a turbojet engine, the fuel is mixed with the air in the combustion chamber position after the compressor. To improve the eciency of the combustion process, the air flow is slowed down to subsonic speeds to improve the mixing and ignition. The Concorde is the only civil airliner to have been equipped with a reheat system (after burner) which would provide up to 26 kN of thrust. The reheat was required for the take-o phase and for a short 10-minute acceleration phase to push the aircraft from Mach 1 to Mach 1.7. The reheat was the primary source of noise in the Concorde. The Concorde was capable of a cruise speed of Mach 2.02 at an altitude of 18 km with a range of 4500 nautical miles. The peculiarity of its turbojet was that the same engine was used to take it from take-o to cruise speeds and back to landing. However, the engine emitting large amount of noise and pollutants due to its poor eciency was the reason for its retirement [3]. 1.3. Mach 3.0 The SR-71 Blackbird, equipped with two J-58JT11D-20 turbo-ramjets, was manufactured by Pratt and Whitney. Each engine produced 145 kN of thrust with the reheat system on and 110 kN without it. The engine consists of nine compression stages fitted on a single spool to the turbine. Below Mach 2, the engine would behave like any turbojet engine. However, at speeds above Mach 2.2, it bypasses bleed air from the compressor straight into the afterburner, which makes the engine behave like a ramjet. Eectively the afterburner would become the combustion chamber. The eect was to improve the eciency of the engine by letting the forward motion of the aircraft produce most of the compression. Similar to a turbojet engine, the combustion process in a ramjet still occurs at subsonic velocity; however, a much larger amount of air is being accelerated due to the series of air bleeds, allowing for greater speeds. The ramjet cannot be started at idle speed like a turbojet because of the required forward motion. Its maximum operating velocity is Mach 6. The engines ran o a new type of jet fuel, JP-7. This fuel having a very high ignition temperature, required Triethylborane (TEB) for the engine and after burner to start. Once ignited, the engine could maintain combustion due to the high service temperature inside the engine. TEB is a hazardous chemical so the refueling process would require specially trained crew and equipment. The Blackbird was capable of speeds of Mach 3.2 at an altitude of 26 km with a range of 3200 nautical miles without refueling [6]. Aerospace 2019, 6, 135 3 of 18 1.4. Beyond Mach 3.0 The NASA Hyper-X program created the X-43 unmanned hypersonic aircraft. This experimental aircraft’s purpose was to validate that an “air breathing scramjet engine” could be used to propel an aircraft to hypersonic speed. The X-43 reached a velocity of Mach 9.6 [7]. A scramjet has no moving parts and does not need to be axially symmetrical like a conventional turbojet engine. The flow throughout the engine remains supersonic, and thus the flow is not slowed down. The theoretical maximum speed using hydrogen fuel is Mach 20. The by-product of this fuel being only water (H O), would provide very clean travel. However, this type of engine required a minimum speed of Mach 6 to function. The forward motion is required to compress the flow to the required pressure and temperature to allow the hydrogen to ignite. The Hypersonic Technology Vehicle 2, HTV-2 for short, was a crewless vehicle which reached Mach 20. As this experimental aircraft was part of the Defense Advanced Research Project Agency (DAPRA) of the United States government, very little information is available in open access. However, the purpose of this experiment was to pave the way of hypersonic engineering and increase the knowledge of the technical challenges to make hypersonic flight reality. The aircraft was launched using a rocket to bring it to its operating altitude and starting velocity above the Pacific Ocean. The drag generated at Mach 20 is such that the aircraft must fly at a very high altitude where the atmosphere is thin enough to reduce the induced drag and to prevent the friction on the body of the aircraft from generating excessive heat. The aircraft was powered using either solid or liquid fuel rocket technology [8]. The aircraft must carry the oxidizer as well, since the amount of oxygen in the atmosphere at high altitudes is insucient for sustainable combustion. This type of propulsion method generates large amount of pollutants, and the corrosive nature of the fuel makes it unsuitable for refueling at airports. 1.5. Future Supersonic Travel The aviation industry has long teased about future technology, which will “soon” be available for commercial aircrafts, new ways to increase the cruise speeds of aircrafts while still reducing the environmental impact all at a reasonable cost. One subsonic project such as the electric turbofan engine partnership between Airbus, Rolls-Royce and Siemens called E-Fan X Hybrid promises greener travel [9]. British aerospace manufacturer, Reaction Engines Limited is working on a precooled heat exchanger engine which will equip supersonic long-distance passenger aircraft [10]. The long-awaited big brother of the SR-71 Blackbird, the SR-72, will be equipped with an over-under configuration turbojet-ramjet combustion design capable of Mach 6 but will remain an unmanned remotely controlled aircraft for military use [11]. The goal of FAME is to design an advanced engine configuration for sustained commercial hypersonic flight. The study will focus on the propulsion system design which will have to meet strict requirements including weight, noise and fuel consumption limits. The study will investigate the improved designs of the inlet and nozzle for both the ramjet and scramjet, whilst also investigating the feasibility of the use of a Dual-Mode Free-Jet combustion chamber design proposed by Trefny and Dippold [12], which is the focus of this report. Combining both would allow for space and weight saving aboard the aircraft while reducing maintenance cost. The dual-mode combustion engine or dual ramjet-scramjet is a system that is a ramjet at low supersonic speeds and then transforms to scramjet at higher Mach number, i.e., 6 to 8. This design also benefits from being useable over a wide range of velocities while keeping reasonable performances. The 2010 numerical investigation by Trefny and Dippold III [12] was limited to hydrocarbon fuel operation over a range of velocities where the geometrical dimensions were slightly altered depending on the flight velocity. Challenges were encountered with extreme static temperatures within the combustion chambers. This report will adapt the Dual-Mode Free-Jet design to fit the FAME propulsion requirements at cruise velocity Mach 8 while using hydrogen instead of hydrocarbon-based fuels. The study will also investigate how eectively altering the fuel to air ratio addresses the excessive temperatures within Aerospace 2019, 6, 135 4 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 4 of 18 same project by Neill and Pesyridis [13]. The FAME aircraft requirements have been previously the combustion chamber. In doing so it extends the work originally conducted as part of the same described by Alkaya et al. [14]. The individual component analyses on which this work is based, for project by Neill and Pesyridis [13]. The FAME aircraft requirements have been previously described by ramjet and scramjet compression systems have been reported in [15–17] and for nozzle performance Alkaya et al. [14]. The individual component analyses on which this work is based, for ramjet and in [18,19]. scramjet compression systems have been reported in [15–17] and for nozzle performance in [18,19]. 2. Dual-Mode Combustion Engines 2. Dual-Mode Combustion Engines As is true for the body of the aircraft, different engine technologies function best at different As is true for the body of the aircraft, dierent engine technologies function best at dierent altitudes and velocities. None of the current commercial aircraft are fitted with engines capable of altitudes and velocities. None of the current commercial aircraft are fitted with engines capable of propelling them to velocities beyond Mach 0.85. This second challenge will be addressed by using propelling them to velocities beyond Mach 0.85. This second challenge will be addressed by using different engine technologies, ramjet and scramjet. However, for the aircraft to be a viable option for dierent engine technologies, ramjet and scramjet. However, for the aircraft to be a viable option for airline companies, it must take off like any other aircraft from a conventional runway. This is airline companies, it must take o like any other aircraft from a conventional runway. This is unfeasible unfeasible with ramjet and scramjet engines since they cannot produce static thrust. To do so, the with ramjet and scramjet engines since they cannot produce static thrust. To do so, the aircraft will aircraft will be equipped with a combination of engines. Turbojet engines deliver thrust by burning be equipped with a combination of engines. Turbojet engines deliver thrust by burning jet fuel with jet fuel with compressed air. The air is compressed through an inlet, then typically through a compressed air. The air is compressed through an inlet, then typically through a multistage compressor multistage compressor before entering the combustion chamber. Once mixed, the air and fuel are before entering the combustion chamber. Once mixed, the air and fuel are ignited to produce thrust. ignited to produce thrust. The hot burnt gases are passed through a turbine and then a nozzle. The The hot burnt gases are passed through a turbine and then a nozzle. The mechanical work produced mechanical work produced by the turbine drives the compressor. Where turbofans are limited to by the turbine drives the compressor. Where turbofans are limited to Mach 0.85, turbojets, like the Mach 0.85, turbojets, like the ones fitted on the Concorde and the SR-71 can operate at higher ones fitted on the Concorde and the SR-71 can operate at higher velocities. Turbojet will be used velocities. Turbojet will be used during taxing and take-off altitude of 0 m up to 10,000 m where it during taxing and take-o altitude of 0 m up to 10,000 m where it will be propelled to a velocity of will be propelled to a velocity of Mach 3.0–3.5. Figure 1 describes the typical flight profile of the plane Mach 3.0–3.5. Figure 1 describes the typical flight profile of the plane from take-o through landing from take-off through landing (13). It shows which engine will be used for each portion of the journey (13). It shows which engine will be used for each portion of the journey including the altitude and including the altitude and speed range. The flight profile includes a stepped approach from the take- speed range. The flight profile includes a stepped approach from the take-o phase to cruise and from off phase to cruise and from cruise to landing. This is needed to allow the pilot to switch between the cruise to landing. This is needed to allow the pilot to switch between the dierent engines, to go from different engines, to go from turbojet to turbo-ramjet, turbo-ramjet to ramjet, ramjet to scramjet and turbojet to turbo-ramjet, turbo-ramjet to ramjet, ramjet to scramjet and the inverse of this process for the inverse of this process for the landing phase. the landing phase. Figure 1. Flight envelope for typical commercial aircraft (13). Figure 1. Flight envelope for typical commercial aircraft (13). 2.1. Dual-Mode Engine Configuration 2.1. Dual-Mode Engine Configuration In this design, a dual-mode combustion chamber which can be used as both ramjet and scramjet is employed. The engine configuration includes, two turbojet engines with two ramjet inlets to channel In this design, a dual-mode combustion chamber which can be used as both ramjet and scramjet is theemployed air into the . The eng turbojetine engines. config When, uration the include aircraft s, tw haso turbojet engi reached a velocit nes wi y of Mach th tw 3o ra andmjet an altitude inlets to of 10,000 m, part of the air is channeled into the dual-mode combustion chamber which is configured in channel the air into the turbojet engines. When, the aircraft has reached a velocity of Mach 3 and an altitude of 10,000 m, part of the air is channeled into the dual-mode combustion chamber which is Aerospace 2020, 7, x FOR PEER REVIEW 5 of 18 Aerospace 2019, 6, 135 5 of 18 configured in ramjet operation. The turbojets and ramjet work hand in hand until Mach 4 is reached at an altitude of 20,000 m. At this stage the turbojet ceases to work. All the air entering the inlets is ramjet operation. The turbojets and ramjet work hand in hand until Mach 4 is reached at an altitude of routed into the ramjet. Finally, once a speed of Mach 6 is reached, the dual-mode combustion chamber 20,000 m. At this stage the turbojet ceases to work. All the air entering the inlets is routed into the is switched from ramjet to scramjet mode. Scramjet mode will remain in operation for the most part ramjet. Finally, once a speed of Mach 6 is reached, the dual-mode combustion chamber is switched for the flight where it will reach its maximum velocity of Mach 8 at altitude 30,000 m. from ramjet to scramjet mode. Scramjet mode will remain in operation for the most part for the flight Another improvement to the body from the previous work [13] is the use of a series of “flaps” where it will reach its maximum velocity of Mach 8 at altitude 30,000 m. or “doors” used to channel the airflow into the different engines. These are shown as dotted lines in Another improvement to the body from the previous work [13] is the use of a series of “flaps” Figure 2. These devices prevent unnecessary drag formation by letting air enter engines which are or “doors” used to channel the airflow into the dierent engines. These are shown as dotted lines in not in use. For example, at Mach 8, the ramjet and turbojet inlets are not being used and thus can be Figure 2. These devices prevent unnecessary drag formation by letting air enter engines which are not “closed off” to make the body of the aircraft more aerodynamically streamlined. The need of two in use. For example, at Mach 8, the ramjet and turbojet inlets are not being used and thus can be “closed turbojet engines is required for safety reasons. In the situation where one turbojet engine fails at low o” to make the body of the aircraft more aerodynamically streamlined. The need of two turbojet altitude, i.e., take-off and landing phase, a second engine is required to ensure the safe landing of the engines is required for safety reasons. In the situation where one turbojet engine fails at low altitude, aircraft. i.e., take-o and landing phase, a second engine is required to ensure the safe landing of the aircraft. Figure 2. Engine configuration. Figure 2. Engine configuration. 2.2. Dual-Mode Combustion Chamber Operating Principles 2.2. Dual-Mode Combustion Chamber Operating Principles The design of the combustion chamber allows for the combustion chamber to be used at either The design of the combustion chamber allows for the combustion chamber to be used at either subsonic or supersonic combustion. This characteristic makes it suitable for a wide range of operating subsonic or supersonic combustion. This characteristic makes it suitable for a wide range of operating velocities. The dual-mode supersonic combustion engine design is patented by Currant and Stull [20]. velocities. The dual-mode supersonic combustion engine design is patented by Currant and Stull [20]. During the use of ramjet, the airflow needs two throats; the first throat is to reduce the flow speed to During the use of ramjet, the airflow needs two throats; the first throat is to reduce the flow speed to subsonic level for combustion, and the second one is for supersonic acceleration in the nozzle (Figure 3). subsonic level for combustion, and the second one is for supersonic acceleration in the nozzle (Figure 3). However, for scramjet the throats are not needed as it can be seen in Figure 4. However, for scramjet the throats are not needed as it can be seen in Figure 4. Aerospace 2020, 7, x FOR PEER REVIEW 6 of 18 Aerospace 2019, 6, 135 6 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 6 of 18 Figure 3. Free-jet dual-mode combustor in subsonic mode. Figure 3. Free-jet dual-mode combustor in subsonic mode. Figure 3. Free-jet dual-mode combustor in subsonic mode. Figure 4. Free-jet dual-mode combustor in supersonic mode. Figure 4. Free-jet dual-mode combustor in supersonic mode. The dierence between a ramjet and scramjet engine is their operation speed range and their Figure 4. Free-jet dual-mode combustor in supersonic mode. The difference between a ramjet and scramjet engine is their operation speed range and their combustion process velocities. The operation speed range of a ramjet lies between Mach 2 and 5 while the combustion p combustion rocess veloc process occurs ities. The oper at subsonic ation speed velocity r . a A nge of scramjet a ramje can t lie function s between from Mach Mach 2 and 4 upwar 5 while ds The difference between a ramjet and scramjet engine is their operation speed range and their the combustion process occurs at subsonic velocity. A scramjet can function from Mach 4 upwards while the combustion process occurs at lower supersonic velocity. combustion process velocities. The operation speed range of a ramjet lies between Mach 2 and 5 while while Ther the co e armbust e technical ion pro challenges cess occu with rs atthis lower type supe of hybrid rsonic veloc engine ity configuration, . due to the modulation the combustion process occurs at subsonic velocity. A scramjet can function from Mach 4 upwards There are technical challenges with this type of hybrid engine configuration, due to the of the thermal location, fuel distribution, flame holding and the actual ignition of the fuel in the while the combustion process occurs at lower supersonic velocity. combustion modulation of chamber the thermal in such locati a on, large fuel cr di oss-section stribution, [fl 13 am ]. e hol When dinused g and the at high actual operating ignition of velocities, the fuel There are technical challenges with this type of hybrid engine configuration, due to the in the combustion chamber in such a large cross-section [13]. When used at high operating velocities, Mach 5 and above, the supersonic combustion occurs in an unconfined supersonic free-jet surrounded modulation of the thermal location, fuel distribution, flame holding and the actual ignition of the fuel by Mach a subsonic 5 and above combustion. , the superson The free-jet ic iscombustion channeled towar occurs in ds a variable an uncon thrf oat ined ar superson ea nozzle.ic Asfree see-jet in in the combustion chamber in such a large cross-section [13]. When used at high operating velocities, surrounded by a subsonic combustion. The free-jet is channeled towards a variable throat area nozzle. Figure 4, the supersonic propulsive stream is not in contact with the wall of the combustion chamber. Mach 5 and above, the supersonic combustion occurs in an unconfined supersonic free-jet T As o adjust see in F the igu nre ozzle 4, the thr sup oatecr rsonic oss section propul ar sive ea, stre a contr am is not ol system in cont is needed act with th to manage e wall of the the com reattachment bustion surrounded by a subsonic combustion. The free-jet is channeled towards a variable throat area nozzle. chamber. To adjust the nozzle throat cross section area, a control system is needed to manage the of the flow when in scramjet operation. Thus, the use of a dual-mode engine could be beneficial as As see in Figure 4, the supersonic propulsive stream is not in contact with the wall of the combustion it reattachment oers up a way of the flow w of building hen in up to scr the amjet operat required Mach ion. Thus, the 8 speed us value e of thr a d ough ual-mode eng an integration ine coof uld two be chamber. To adjust the nozzle throat cross section area, a control system is needed to manage the beneficial as it offers up a way of building up to the required Mach 8 speed value through an engine systems. This leads to space, weight and maintenance cost savings. Having less combustion reattachment of the flow when in scramjet operation. Thus, the use of a dual-mode engine could be chambers integration of incr teases wo engine the space systems. Th for extra is lpayload eads to spac or fuel e, weight and r educes and maint theen need ance to cost maintain savingsmultiple . Having beneficial as it offers up a way of building up to the required Mach 8 speed value through an less combustion chambers increases the space for extra payload or fuel and reduces the need to combustion chambers. The localized heating eects are negated due to the flow not being in contact integration of two engine systems. This leads to space, weight and maintenance cost savings. Having maintain multiple combustion chambers. The localized heating effects are negated due to the flow with the combustion chamber walls. A secondary flow of air can be introduced into the engine to act less combustion chambers increases the space for extra payload or fuel and reduces the need to not being in contact with the combustion chamber walls. A secondary flow of air can be introduced as a cooling agent whilst increasing the pressure. However, this consideration will not be studied in maintain multiple combustion chambers. The localized heating effects are negated due to the flow into the engine to act as a cooling agent whilst increasing the pressure. However, this consideration this paper. not being in contact with the combustion chamber walls. A secondary flow of air can be introduced will not be studied in this paper. The high-speed air enters the engine through the inlet where it is compressed through a series of into the engine to act as a cooling agent whilst increasing the pressure. However, this consideration The high-speed air enters the engine through the inlet where it is compressed through a series shockwaves. As it is compressed, it is slowed down to subsonic speeds in the case of a ramjet engine and will not be studied in this paper. of shockwaves. As it is compressed, it is slowed down to subsonic speeds in the case of a ramjet remains supersonic in scramjet engine. However, the flow is still slowed down to improve the eciency The high-speed air enters the engine through the inlet where it is compressed through a series engine and remains supersonic in scramjet engine. However, the flow is still slowed down to improve of the combustion chamber. Once in the combustion chamber, the fuel is injected and mixed with the of shockwaves. As it is compressed, it is slowed down to subsonic speeds in the case of a ramjet the efficiency of the combustion chamber. Once in the combustion chamber, the fuel is injected and air. In the case of ramjet, flame holders (igniters) may be required to maintain a steady combustion of engine and remains supersonic in scramjet engine. However, the flow is still slowed down to improve mixed with the air. In the case of ramjet, flame holders (igniters) may be required to maintain a steady the fuel. In scramjet engine, the fuel automatically ignites due to the extreme temperatures of the flow the efficiency of the combustion chamber. Once in the combustion chamber, the fuel is injected and combustion of the fuel. In scramjet engine, the fuel automatically ignites due to the extreme entering the combustion chamber. Finally, for both engines, the flow is expanded and ejected though mixed with the air. In the case of ramjet, flame holders (igniters) may be required to maintain a steady temperatures of the flow entering the combustion chamber. Finally, for both engines, the flow is the nozzle. The nozzle’s role is to convert the pressure and thermal energy of the air and fuel into combustion of the fuel. In scramjet engine, the fuel automatically ignites due to the extreme kinetic energy which in turn generates thrust. temperatures of the flow entering the combustion chamber. Finally, for both engines, the flow is Aerospace 2019, 6, 135 7 of 18 Both the ramjet and the scramjet use the forward motion of the engine to compress the incoming air. This means no moving parts are required to compress the air to the required pressures and temperatures for the combustion process. The Dual-Mode Free-Jet combustor has only been tested in a simulated computational environment thus far. This paper will validate the result using CFD analysis. 2.3. Fuel The Dual-Mode Free-Jet combustor by Triny and Dupold [13] was modelled using ethylene fuel. This paper will be carrying out simulations using hydrogen as a fuel; hence, this section will compare the characteristics of ethylene and hydrogen fuel. Fuel properties are provided in Table 1, below. Table 1. Comparison of fuel properties. Fuel Energy/Mass Energy/Volume Density Hydrogen 116.7 8.2 71 Methane 50.2 20.8 424 Ethylene 47.2 26.8 568 A study from the University of Queensland [21] investigates the suitability of hydrogen, methane and ethylene fuels at Mach 8. From this analysis, it can be clearly seen that despite hydrogen having the higher energy content per mass, it is also the least dense of the three fuels. Therefore, consideration for the need of large storage tanks will have to be made. This is one of the driving factors for using a dual-mode combustor, the space gained by combining two separated combustion chambers into one, can be allocated to increase fuel storage. Another influencing factor for choosing pure hydrogen fuel as opposed to hydrocarbon-based fuels, is the sustainability it oers. Its only by-product is water vapor (H O). The high energy content of hydrogen provides a significant advantage as highlighted by El-Sayed [17]. The specific impulse potential of hydrogen is significantly higher allowing for better performing accelerations. More so, it can be used to fuel turbojets, ramjets and scramjet engines; hence, it could be used from take-o to cruise velocity, therefore relieving the need for alternative fuel for the other engines. The water vapour produced by the combustion of hydrogen does have a polluting eect at very high altitude [22]. The presence of water vapor in the stratosphere, has a warming eect of the troposphere, hence being a contributor in global warming [23]. 2.4. Injection and Combustion The challenge with supersonic combustion is achieving sucient mixing of the fuel and air within the combustor [24]. Due to the high flow velocities, the fuel has only a short amount of time to mix [25]. The extreme conditions must allow the fuel to atomize, mix, ignite and produce a steady and reliable combustion. This is especially challenging for hydrocarbon fuels, due to the slow reaction rates and the short residence time involved that limits fuel-air mixing [26]. Taking into consideration the Rayleigh flow eect, heat addition from adding fuel into the supersonic flow may result in a decrease of the stagnation pressure. An excessive decrease in pressure will impact the eciency of the engine negatively and may result in the supersonic flow slowing down to subsonic speeds. Hence, for the fuel to add energy to the flow, the air must not enter the combustion chamber at too high a temperature. The injection system used by Trefny and Dippold III [13] consists of 3 annular ring injectors. For any given amount of fuel, there is a finite amount of oxygen required to burn it. Excess in oxygen does not hinder the combustion, and when complete combustion is preferred, oxygen is typically supplied in excess. The chemical interaction between the fuel and the oxidizers such as ethylene-air and nitrogen, and hydrogen-air and nitrogen, creates complex reactions species. The combustion simulation must model these sequences of molecular reactions, which release energy while being converted into products. Aerospace 2020, 7, x FOR PEER REVIEW 8 of 18 ethylene-air and nitrogen, and hydrogen-air and nitrogen, creates complex reactions species. The combustion simulation must model these sequences of molecular reactions, which release energy Aerospace 2019, 6, 135 8 of 18 while being converted into products. 3. Computational Methodology 3. Computational Methodology The computational simulations have the goal to simulate and quantify the performances of the The computational simulations have the goal to simulate and quantify the performances of the adapted Dual-Mode Free-Jet design to fit the propulsion requirements at cruise velocity Mach 8 while adapted Dual-Mode Free-Jet design to fit the propulsion requirements at cruise velocity Mach 8 while using hydrogen instead of hydrocarbon-based fuels. Commercial software Ansys Fluent was used to using hydrogen instead of hydrocarbon-based fuels. Commercial software Ansys Fluent was used to perform the simulations. Investigation on the effect adjusting the fuel to air ratio, from equivalence perform the simulations. Investigation on the eect adjusting the fuel to air ratio, from equivalence ratio 0.4 to 1.0, has on the temperatures within the combustion chamber will also be studied. ratio 0.4 to 1.0, has on the temperatures within the combustion chamber will also be studied. 3.1. 3.1. Geo Geometry metry The The geometry geometry of of the comb the combustor ustor, used , used for for all all c cases, ases,was was identical identical to the one to the one used usedin in the the 2010 2010 study study by by Trefny an Trefny andd D Dippold ippold III [13] (Figure III [13] (Figure 55 and and T Table able 2). 2). The dimen The dimensions sions were were kept kept identic identical al for for the the validation validation case and we case and werere modified modified to suit the to suit the FAME FAM objectives. E objectives. T The dotted he dotted lin line between e between A and K is A and K is the symmetry the symmetry line. The line. The step step at D and C is at D and C is to to induce induce flow flow separation separation from from the wall. The curve through the wall. The curve through point point G– G–H–I H–I is isdefined defined with two ta with two tangent ngent constra constraints ints to the nei to the neighboring ghboring li line ne F– F–G G a and nd I– I–JJ and and f fixed ixed height height for for t the he t thr hroat oat, , point point H. The n H. The nozzle ozzle sect section ion is is not not papart rt of the combustor fi of the combustor nal desi final design, gn, but ibut ts only its purpose is to demonstrate expansion of the flow after passing the throat (H) during simulation. The only purpose is to demonstrate expansion of the flow after passing the throat (H) during simulation. The ramjet and ramjet and scramjet nozzle scramjet nozzles s designed designed are dimensione are dimensioned d to be fit to be tfitted ed at point at point H. T H.h The e inject injection ion sysystem stem is positioned between points A and B, as shown in Figure 6. is positioned between points A and B, as shown in Figure 6. Figure Figure 5. 5. Combus Combustion tion c chamber hamber geometry. geometry. Table 2. Combustion chamber and injector dimensions for ethylene-based engine design. Table 2. Combustion chamber and injector dimensions for ethylene-based engine design. Combustion Chamber Dimensions Injector Dimensions Combustion Chamber Dimensions Injector Dimensions Point X (mm) Y (mm) Point X (mm) Y (mm) Point X (mm) Y (mm) Point X(mm) Y(mm) A 0.0 0.0 L 0.0 17.20 A 0.0 0.0 L 0.0 17.20 B 0.0 119.5 M 0.0 22.60 B 0.0 119.5 M 0.0 22.60 C 111.0 131.4 N 0.0 57.0 C 111.0 131.4 N 0.0 57.0 D 111.0 131.4 O 0.0 62.40 D 111.0 131.4 O 0.0 62.40 E 595.2 351.0 P 0.0 96.80 F E 59 1297.1 5.2 351.0351.0 QP 0.0 0.0 96.80 102.20 G 1407.5 248.2 F 1297.1 351.0 Q 0.0 102.20 H 1533.1 196.7 G 1407.5 248.2 I 1622.6 216.0 H 1533.1 196.7 J 2200.0 500.0 I 1622.6 216.0 K 2200.0 0.0 J 2200.0 500.0 K 2200.0 0.0 Aerospace 2019, 6, 135 9 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 9 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 9 of 18 Figure 6. Close-up view of the injection system. Figure 6. Close-up view of the injection system. Figure 6. Close-up view of the injection system. The dimension of the injectors had to be drastically reduced in size to accommodate the The dimen The dimension sion of the of the injectors injectors had t had to be drastically o be drastic reduced ally red in size uced to in size accommodate to accthe ommodate t replacement he replacement fuel, from ethylene to hydrogen due to the higher calorific content. In a ramjet-scramjet replac fuel, fr ement fuel, from eth om ethylene to hydr ylene to hydr ogen due to ogen due to the higherthe higher c calorific content. alorific content. In a ramjet-scramjet In a ramjet-scramjet engine, engine, the combustion chamber can be asymmetrically shaped, meaning it does not need to be engine, the c the combustion ombustion c chamber can hamber c be asymmetrically an be asymmetr shaped, ically sh meaning aped, mean it doesing notit need does not nee to be cylindrical d to be cylindrical or axially symmetrical [22]. Figure 7 shows the asymmetrical design of the Dual-Mode cylin or axially drical or symmetrical axially sym [22 m ].etrica Figur l [2 e2] 7. Fi shows gure the 7 sho asymmetrical ws the asymdesign metricaof l de the sign of Dual-Mode the Dual Fr -Mode ee-Jet Free-Jet combustion chamber. Free-J combustion et comb chamber ustion cha . mber. Figure 7. 3D view of dual-mode rectangular combustion chamber. Figure 7. 3D view of dual-mode rectangular combustion chamber. Figure 7. 3D view of dual-mode rectangular combustion chamber. 3.2. Boundary Types 3.2. Boundary Types The inlet (fuel and air) and outlet boundaries were specified using pressure. The computational 3.2. Boundary Types domain The inlet (fue was halved l and employing air) and outlet bound symmetry conditions aries were sp soecif asied usin to reduce g pressu the computational re. The computat e ion ort. al The inlet (fuel and air) and outlet boundaries were specified using pressure. The computational The boundary conditions are described in Table 3. domain was halved employing symmetry conditions so as to reduce the computational effort. The domain was halved employing symmetry conditions so as to reduce the computational effort. The boundary conditions are described in Table 3. boundary conditions are described in Table 3. Table 3. Boundary conditions. Table 3. Boundary conditions. Total Pressure Table 3. (Pa) Boundary conditions Temperature . (K) Species Total Pressure (Pa) Temperature (K) Species Inlet 150,105 3012 O —21.15% Total Pressure (Pa) Temperature (K) Species Injector 150,105 547 C H —100% Inlet 150,105 3012 O2—21.15% 2 4 Inlet 150,105 3012 O2—21.15% Outlet 1580 227 Injector 150,105 547 C2H4—100% Injector 150,105 547 C2H4—100% Outlet 1580 227 Outlet 1580 227 3.3. Numerical Model 3.3. Numerical Model There is no single turbulence model able to universally predict every turbulent flow. Choosing the 3.3. Numerical Model correct model can have significant impact on the accuracy of the flow modelling. The calculation carried There is no single turbulence model able to universally predict every turbulent flow. Choosing There is no single turbulence model able to universally predict every turbulent flow. Choosing out by NASA used a time-accurate, fully implicit code developed in-house [13]. In this study the k-" the correct model can have significant impact on the accuracy of the flow modelling. The calculation the correct model can have significant impact on the accuracy of the flow modelling. The calculation model was employed. RANS models are applicable to a wide range of applications while proving carried out by NASA used a time-accurate, fully implicit code developed in-house [13]. In this study carried out by NASA used a time-accurate, fully implicit code developed in-house [13]. In this study appropriately accurate for this study. Density based models are recommended for supersonic simulation the k-ε model was employed. RANS models are applicable to a wide range of applications while the k-ε model was employed. RANS models are applicable to a wide range of applications while proving appropriately accurate for this study. Density based models are recommended for proving appropriately accurate for this study. Density based models are recommended for Aerospace 2020, 7, x FOR PEER REVIEW 10 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 10 of 18 Aerospace 2019, 6, 135 10 of 18 supersonic simulation since they provide good convergence properties as opposed to pressure-based supersonic simulation since they provide good convergence properties as opposed to pressure-based since they provide good convergence properties as opposed to pressure-based models. For the solution models. For the solution of the governing equations, an implicit approach with the Roe-flux models. For the solution of the governing equations, an implicit approach with the Roe-flux of the governing equations, an implicit approach with the Roe-flux dierence splitting for calculating difference splitting for calculating the convective flux was used. The flow, turbulent kinetic energy, difference splitting for calculating the convective flux was used. The flow, turbulent kinetic energy, the convective flux was used. The flow, turbulent kinetic energy, and turbulent dissipation were set as and turbulent dissipation were set as second order upwind to obtain more accurate solutions. Single and turbulent dissipation were set as second order upwind to obtain more accurate solutions. Single second order upwind to obtain more accurate solutions. Single step, volumetric reactions and finite step, volumetric reactions and finite rate reaction for the turbulence-chemistry interaction were step, volumetric reactions and finite rate reaction for the turbulence-chemistry interaction were rate reaction for the turbulence-chemistry interaction were enabled with the hydrogen–air mixture enabled with the hydrogen–air mixture (H2 + O2), to model the scramjet combustion. The Y+ value enabled with the hydrogen–air mixture (H2 + O2), to model the scramjet combustion. The Y+ value (H + O ), to model the scramjet combustion. The Y+ value was maintained less than 1. A Species was maintained less than 1. A Species Transport Laminar Finite Rate model will be employed to was 2 maint 2 ained less than 1. A Species Transport Laminar Finite Rate model will be employed to T model ransport the vol Laminar umetri Finite c reaRate ction combustion process model will be employed where to model the turb the ulent fluctuat volumetric reaction ions are igno combustion red. It model the volumetric reaction combustion process where the turbulent fluctuations are ignored. It process where the turbulent fluctuations are ignored. It uses the Arrhenius Rate of Reaction Law. uses the Arrhenius Rate of Reaction Law. The chemical mechanisms are integrated into Fluent uses the Arrhenius Rate of Reaction Law. The chemical mechanisms are integrated into Fluent The ANSY chemical S using t mechanisms he CHEMKIN t are integrated ool. This so intoft Fluent ware ANSYS is used for usingso the lving compl CHEMKIN ex chemica tool. Thisl softwar kinetices ANSYS using the CHEMKIN tool. This software is used for solving complex chemical kinetics is used for solving complex chemical kinetics problems. problems. problems. 3.4. Mesh 3.4. Mesh 3.4. Mesh The computational domain was meshed using a quadrilateral mesh and using edge sizing The computational domain was meshed using a quadrilateral mesh and using edge sizing The computational domain was meshed using a quadrilateral mesh and using edge sizing wherever possible. Figure 8 shows the grid pattern generated to the NASA specifications. The grid wherever possible. Figure 8 shows the grid pattern generated to the NASA specifications. The grid wherever possible. Figure 8 shows the grid pattern generated to the NASA specifications. The grid shown in the figure is with reduced mesh density for better clarity. The node number used for shown in the figure is with reduced mesh density for better clarity. The node number used for shown in the figure is with reduced mesh density for better clarity. The node number used for validation was in accordance with number used by NASA, approximately 392,000. validation was in accordance with number used by NASA, approximately 392,000. validation was in accordance with number used by NASA, approximately 392,000. Figure 8. Computational mesh. Figure 8. Figure 8. Compu Computational tational mes mesh. h. The grid independence study was carried out using ANSYS Workbench, where 5 dierent mesh The grid independence study was carried out using ANSYS Workbench, where 5 different mesh The grid independence study was carried out using ANSYS Workbench, where 5 different mesh densi densities ties were crea were created, ted, a a 50,000, 50,000, 70 70,000, ,000, 110,000, 110,000, 177,000 177,000 a an nd d 267 267,000. ,000. The The 11 110-k 0-k cel cell l mesh showed the mesh showed the densities were created, a 50,000, 70,000, 110,000, 177,000 and 267,000. The 110-k cell mesh showed the fastest convergence to steady mass flow rate of 0 kg/s (Figure 9). Coarser meshes provided a very fastest convergence to steady mass flow rate of 0 kg/s (Figure 9). Coarser meshes provided a very fastest convergence to steady mass flow rate of 0 kg/s (Figure 9). Coarser meshes provided a very unstable v unstable value alue at the be at the beginning ginning o of f the the simulation simulation a and nd wou would ld e end nd up ta up taking king a a gr grea eater ter a amount mount of of ti time me unstable value at the beginning of the simulation and would end up taking a greater amount of time to converge. Denser meshes provided more stable results; however, the convergence would require to converge. Denser meshes provided more stable results; however, the convergence would require to converge. Denser meshes provided more stable results; however, the convergence would require a gre a greater ater amount amount of time to converge, particular of time to converge, particularlyly for fo267,000 r 267,000 ce cells lls and and ab above. ove It.was It was dec decided ided upon upon the a greater amount of time to converge, particularly for 267,000 cells and above. It was decided upon results of the mesh refinement study to carry all simulations with a mesh count of 110,000 cells. the results of the mesh refinement study to carry all simulations with a mesh count of 110,000 cells. the results of the mesh refinement study to carry all simulations with a mesh count of 110,000 cells. Figure 9. Mesh refinement study. Figure 9. Mesh refinement study. Figure 9. Mesh refinement study. Aerospace 2019, 6, 135 11 of 18 3.5. Species Model The three-step reaction with 7 species for ethylene–air is described in Table 4 [23] and the one step reaction for hydrogen–oxygen is described in Table 5. Table 4. Ethylene–air chemical reaction and products. Reactions Products Species Stoich Coecient Rate Exponent Species Stoich Coecient Rate Exponent C H 1 1 CO 2 0 2 4 O 1 1 H 2 0 2 2 Pre exponential factor Activation energy Temperature exponent 11 8 2.10 10 1.50 10 0 Species Stoich Coecient Rate Exponent Species Stoich Coecient Rate Exponent CO 2 1 CO 2 0 O 1 1 Pre exponential factor Activation energy Temperature exponent 5 7 3.48 10 8.43 10 2 Species Stoich Coecient Rate Exponent Species Stoich Coecient Rate Exponent H 2 1 H O 2 0 2 2 O 1 1 Pre exponential factor Activation energy Temperature exponent 3.00 10 0 1 Table 5. Hydrogen–oxygen chemical reaction and products. Reactions Products Species Stoich Coecient Rate Exponent Species Stoich Coecient Rate Exponent H 1 1 H O 1 0 2 2 O 0.5 1 Pre exponential factor Activation energy Temperature exponent 8 7 9.87 10 3.1 10 0 4. Results and Discussion 4.1. Validation In this section the numerical methodology adopted in this study is validated by comparing the simulation results with experimental results provided by NASA [19]. The results in this section will be limited to flow property comparison at Mach 8. The NASA numerical experiment shown on the right in Figure 10 shows fast incoming air entering the combustion chamber whilst the fuel is traveling at a lower speed than Mach 1. The flow continues to slow down until reaches past the throat at about Mach 2.0 where it proceeds to speed up to Mach 3.6 in the nozzle. In the validation case shown on the left, the fuel injection is seen to be similarly slower than the incoming air; however, the flow does not slow down as dramatically as in the NASA study. This may be because the fuel is acting as a speed reducing agent as it mixes with the air. The fuel does not seem to penetrate as far into the combustion chamber. Another dierence can be noticed that the stagnating recirculation zone is separated by an intermediate zone of flow before the free-jet stream of air at Mach 2.4 in the computational analysis. Nonetheless, the flow reattaches in both cases at the throat at a speed above Mach 1 and accelerates downstream. Aerospace 2019, 6, 135 12 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 12 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 12 of 18 (a) (b) (a) (b) Figure 10. Figure 10. Comparison of M Comparison of Maa ch contour; ( ch contour; ( a) com a) com pu ptu ational resu tational resu lt, ( lt b, ( ) exp b) exp erimental result [17]. erimental result [17]. Figure 10. Comparison of Mach contour; (a) computational result, (b) experimental result [17]. The The extreme temperature extreme temperatures s in the recirculation reg in the recirculation region ion ofof 505000 00 kel kelvin vin (9500 (9500 Ranki Rankin) n) were observed were observed The extreme temperatures in the recirculation region of 5000 kelvin (9500 Rankin) were observed in the NASA numerical experiment, Figure 11b. This was successfully replicated in the validation in the NASA numerical experiment, Figure 11b. This was successfully replicated in the validation in the NASA numerical experiment, Figure 11b. This was successfully replicated in the validation study. The subsonic recirculating flow can be seen to be in the region of 5000 kelvin degrees. These study. The subsonic recirculating flow can be seen to be in the region of 5000 kelvin degrees. These high study. The subsonic recirculating flow can be seen to be in the region of 5000 kelvin degrees. These high temperature causes ionizing of atoms, which is undesirable for combustion. This can be temperature causes ionizing of atoms, which is undesirable for combustion. This can be addressed by high temperature causes ionizing of atoms, which is undesirable for combustion. This can be addressed by adjusting the amount of fuel being introduced into the combustion chamber. Although adjusting the amount of fuel being introduced into the combustion chamber. Although dierences exist addressed by adjusting the amount of fuel being introduced into the combustion chamber. Although differences exist in the Mach and temperature plots, the setup provides results, which provide in the Mach and temperature plots, the setup provides results, which provide sucient validations for differences exist in the Mach and temperature plots, the setup provides results, which provide sufficient validations for further analysis. further analysis. sufficient validations for further analysis. (a) (b) (a) (b) Figure 11. Comparison of temperature distribution, (a) computational result, (b) historical numerical Figure Figure 11. 11. Co Comparison mparison of te of temperatur mperature distribution, ( e distribution, (a a) c ) computational omputational resu result, lt, ( (b b) ) historical numerical historical numerical experimental result [17]. experimental experimental result [17] result [17].. 4.2. Effect of Equivalence Ratio on the Flow Characteristics 4.2. Eect of Equivalence Ratio on the Flow Characteristics 4.2. Effect of Equivalence Ratio on the Flow Characteristics Following the validation run of the NASA design, the model was modified to meet the Following the validation run of the NASA design, the model was modified to meet the specifications specifications for the present study. The numerical methodology parameters used are identical to the Following the validation run of the NASA design, the model was modified to meet the for the present study. The numerical methodology parameters used are identical to the validation case, validation case, except that the species and materials parameters have been changed to reflect the use specifications for the present study. The numerical methodology parameters used are identical to the except that the species and materials parameters have been changed to reflect the use of hydrogen fuel. of hydrogen fuel. The effect of equivalence ratio (EQ) on the flow characteristics has been studied in validation case, except that the species and materials parameters have been changed to reflect the use The eect of equivalence ratio (EQ) on the flow characteristics has been studied in detail. detail. of hydrogen fuel. The effect of equivalence ratio (EQ) on the flow characteristics has been studied in detail. 4.2.1. Mach Number 4.2.1. Mach Number 4.2.1. Mach As seen Nin um Figur ber e 12a–d, the Mach velocity at the inlet is below Mach 1.0. This suggests that As seen in Figure 12a–d, the Mach velocity at the inlet is below Mach 1.0. This suggests that a normal shock perpendicular to the flow has formed at the very entrance of the inlet, which causes the a normal shock perpendicular to the flow has formed at the very entrance of the inlet, which causes As seen in Figure 12a–d, the Mach velocity at the inlet is below Mach 1.0. This suggests that a flow to slow down as explained above. This shock formation causes the flow to go into subsonic the flow to slow down as explained above. This shock formation causes the flow to go into subsonic normal shock perpendicular to the flow has formed at the very entrance of the inlet, which causes the velocities in the combustion chamber. In essence, the scramjet engine is behaving as a ramjet where velocities in the combustion chamber. In essence, the scramjet engine is behaving as a ramjet where flow to slow down as explained above. This shock formation causes the flow to go into subsonic combustion occurs at subsonic flows. The free-jet is at subsonic velocity instead of supersonic. The combustion occurs at subsonic flows. The free-jet is at subsonic velocity instead of supersonic. velocities in the combustion chamber. In essence, the scramjet engine is behaving as a ramjet where free-jet is surrounded by a lower speed static region of flow. The fuel can be seen to be traveling The free-jet is surrounded by a lower speed static region of flow. The fuel can be seen to be traveling combustion occurs at subsonic flows. The free-jet is at subsonic velocity instead of supersonic. The through the free-jet at a lower speed and the surrounding air from the inlet. Looking at the Mach through the free-jet at a lower speed and the surrounding air from the inlet. Looking at the Mach free-jet is surrounded by a lower speed static region of flow. The fuel can be seen to be traveling ratio along the axial centerline distance of the combustor (Figure 13), significant fluctuations occur at ratio along the axial centerline distance of the combustor (Figure 13), significant fluctuations occur at through the free-jet at a lower speed and the surrounding air from the inlet. Looking at the Mach Equivalence ratio 0.4 where little fuel is injected. The peak velocity occurs in the divergent part of the Equivalence ratio 0.4 where little fuel is injected. The peak velocity occurs in the divergent part of the ratio along the axial centerline distance of the combustor (Figure 13), significant fluctuations occur at combustor, between 0.5 m and 2.2 m in all cases. However, only for equivalence ratio 0.4 does the combustor, between 0.5 m and 2.2 m in all cases. However, only for equivalence ratio 0.4 does the Equivalence ratio 0.4 where little fuel is injected. The peak velocity occurs in the divergent part of the Mach reaches 1.4 times the inlet speed. Mach reaches 1.4 times the inlet speed. combustor, between 0.5 m and 2.2 m in all cases. However, only for equivalence ratio 0.4 does the Mach reaches 1.4 times the inlet speed. Aerospace 2019, 6, 135 13 of 18 Aerospace Aerospace 2020 2020, , 7 7, x FO , x FOR P R PEER EER RE REVIEW VIEW 13 of 13 of 18 18 ( (a a) ( ) (b b) ) ( (c c) ( ) (d d) ) Figure Figure 12. Figure 12. 12. ( ((a a a) ) ) Mach number Mach number Mach number contour for equivalence ratio of 0. contour for equivalence ratio of 0. contour for equivalence ratio of 0.4, 4, 4, ( ( (b b b) ) ) Mach Mach nu Mach nu number m mb ber contou er contou contourr for r for for equivalence equivalence equivalence ratio of ratio of ratio of 0.6, 0.6 0.6 (c ,, ( ( ) Mach c c) ) Mach nu Mach nu number m mb b contour er contou er contou forr for eq r for eq equivale uivalence rat uivalence rat nce ratio of 0.8, iio o of 0 of 0 (d) Mach ..8, ( 8, (d d) Ma ) Ma number ch ch numbe numbe contourrr for contour for equivalence ratio contour for equivalence ratio equivalence ratio of 1.0. of 1.0. of 1.0. Figure 13. Figure 13. Variation of Variation of Mach r Mach ra atio with tio with ax axial ial distanc distance e.. Figure 13. Variation of Mach ratio with axial distance. The equivalence ratio mixture 0.6 peeks at a ratio 1, meaning its maximum speed is equal to The equiv The equiva ale lence rat nce rati io mixt o mixture ure 0.6 pe 0.6 peeks eks at a r at a ra ati tio o 1, 1, m me eanin aning g it its m s ma aximum ximum speed speed is is e eq qu ua al l t to o it its s its inlet speed (Mi). EQ 0.8 and 1.0 show poor Mach ratios below 1 throughout the combustion inlet inlet sp spee eed ( d (M Mi) i). EQ . EQ 0. 0.8 a 8 an nd 1. d 1.0 show p 0 show poor Mach oor Mach rat rati ios b os be elow low 1 t 1 th hrougho roughout ut t th he com e comb bust usti ion c on ch ham amb be er. r. chamber. The Mach plots in Figure 12a–d shows the flow velocities to be above Mach 1 prior to The Mach plots in The Mach plots in Figure Figure 12a–d 12a–d shows shows the flow the flow velo veloci citi ties to be es to be a ab bove Ma ove Mach ch 1 1 p pr riio or to r to rea reac ching the hing the reaching the throat, hence allowing the flow to accelerate in the nozzle. However, the expected Mach throa throatt, hence al , hence alllo owi wing the fl ng the flow to a ow to accelera ccelerate te iin n the n the no ozzle. How zzle. Howe ever, the ver, the expe expected M cted Ma ach ve ch velocity locity velocity between Mach 2.5 to 3.0 required for the scramjet nozzle to produce enough thrust was b be et tw ween Mac een Mach h 2. 2.5 t 5 to o 3. 3.0 r 0 re equ quired ired for t for th he e scr scra am mjjet et nozzle nozzle to produc to produce enough e enough thr thru ust was not st was not achieved. achieved. not achieved. As the amount of fuel is increased, from equivalence ratio 0.4 to 1.0, the Mach value As the amo As the amou unt of fuel nt of fuel is is incre increa ased sed,, fr from equi om equiv va ale lence rat nce ratiio o 0. 0.4 4 t to o 1. 1.0, t 0, th he e Mach v Mach va allu ue e rem rema ain inss remains unchanged. This suggests that a lean fuel-air mixture would have an as good eect of thrust unchan unchanged. ged. This s This su ugg ggest ests s t th hat at a a le lean f an fu uel el-a -air m ir miixt xtu ur re wou e woulld h d ha av ve e an an as as good good effect effect o off t th hru rust st as as a a as a stoichiometric mixture. stoichiometric mixture. stoichiometric mixture. 4.2.2. Velocity 4. 4.2. 2.2. Ve 2. Velocit locity y Similar to the Mach plots, the velocity plots in Figure 14a–d show a relatively fast subsonic, Si Simila milar to the r to the Ma Mach p ch pl lots, t ots, th he vel e velo oci city plots i ty plots in n Fi Figure gure 1 14a– 4a–d show a rela d show a relati tive vely fa ly fast su st subsoni bsonic c, , free- free- free-jet flow surrounded by a slower, recirculating zone of air above it. The freestream jet reattaches jet flow jet flow surro surrou unded by a slower, rec nded by a slower, reci irc rculating ulating zone of zone of air air above above it. T it. Th he free e freestream stream jet reattach jet reattaches at the es at the at the throat where in reaches a velocity of about 650 to 850 m/s depending on the fuel content of tth hroat roat where where in reach in reache es s a a v ve elocit locity y of of a ab bout out 65 650 0 to 85 to 850 0 m/s dependi m/s dependin ng on the f g on the fu uel el content of content of the the the mixture. The relationship between fuel-to-air ratio and flow acceleration becomes clear in the mixture. The mixture. The relationsh relationship b ip be etween tween fu fuel- el-t to- o-ai air ra r rati tio o a an nd fl d flow a ow accel ccele era rati tion becom on become es cl s clea ear r i in n the nozzl the nozzle e nozzle as shown by Figure 15. With an EQ of 0.4 and 0.6, the lean mix of fuel does not allow the air as as shown b shown by y Fig Figu ure re 1 15 5. W . Wiitth h an EQ an EQ of of 0. 0.4 4 and and 0. 0.6, t 6, th he le e lean m an miix x of f of fu uel el does does not not al allow t low th he e ai air t r to o to accelerate above 960 m/s in the nozzle. However, with higher fuel mixtures, EQ 1.0 and EQ 0.8, accel accele erat rate e ab abov ove 96 e 960 m 0 m/ /s i s in n t th he nozzl e nozzle. e. Howev Howeve er, wit r, with h higher f higher fu ue ell m miixt xtures ures, EQ , EQ 1. 1.0 and EQ 0 and EQ 0 0..8, t 8, th he e the flow accelerates beyond 1000 m/s. Figure 14b,c shows two dark red streaks of fast moving flow in flow flow acc acce eler lera at te es beyon s beyond d 1000 m/s. 1000 m/s. F Fi igu gur re e 14b,c 14b,c show shows s two d two da ark rk re red d stre streaks o aks of f fast moving fast moving flo flow w in in the the the nozzle at EQ 0.8 and EQ 1.0, which are not present at EQ 0.4 and are only faintly visible at EQ 0.6. nozz nozzle at le at EQ 0. EQ 0.8 and 8 and EQ 1. EQ 1.0, 0, wh which are not ich are not p pr resent esent a at t EQ 0. EQ 0.4 4 and and are on are only ly fa faint intlly v y viisib siblle e at at E EQ Q 0. 0.6. 6. This, contrary to the Mach flow plots, would suggest that a stoichiometric mixture would be preferable This, contrary to the Mach flow plots, would suggest that a stoichiometric mixture would be This, contrary to the Mach flow plots, would suggest that a stoichiometric mixture would be to a lean fuel-air mixture. The disparity between the Mach and Velocity plots are due to the local speed prefer preferable to able to a le a lean an fue fuell-air mixture. The disp -air mixture. The dispar arity be ity between the Mach and tween the Mach and Velocity plots ar Velocity plots are d e du ue to e to the local speed of sound changing throughout the combustion chamber caused by the temperature the local speed of sound changing throughout the combustion chamber caused by the temperature incre increa ase, se, a as s s seen in een in t th he nex e next t sect section. ion. Aerospace 2019, 6, 135 14 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 14 of 18 of sound changing throughout the combustion chamber caused by the temperature increase, as seen in Aerospace 2020, 7, x FOR PEER REVIEW 14 of 18 the next section. (a) (b) (a) (b) (c) (d) (c) (d) Figure 14. (a) Velocity contour for equivalence ratio of 0.4, (b) Velocity contour for equivalence ratio of 0.6, (c) Velocity contour for equivalence ratio of 0.8, (d) Velocity contour for equivalence ratio of Figure Figure 14. 14. ( (a a) ) V Velocity contour for equivalence ratio of 0.4, ( elocity contour for equivalence ratio of 0.4, (b) bV ) Velocity contour for equivalence ratio elocity contour for equivalence ratio of 1.0. 0.6, of 0.6, ( (c) V celocity ) Velocity contour fo contour for equivalence r equivalence r ratio a of tio of 0 0.8, (d.8, ( ) Velocity d) Velocity contour for equivalence ratio of contour for equivalence ratio of 1.0. 1.0. Figure 15. Figure 15. Variation of f Variation of flow low velocity velocity with ax with axial ial d distance. istance. Figure 15. Variation of flow velocity with axial distance. 4.2.3. Temperature 4.2.3. Temperature 4.2.3. Temperature As expected, the combustion process releases energy in the form of heat proportionally to the As expected, the combustion process releases energy in the form of heat proportionally to the amount of fuel introduced into the combustion. Figure 16 shows the static temperature of the fuel amount of As expected, fuel introduced i the combustion process re nto the combustileases ene on. Figure rgy i 16 show n the f s t oh rm e stat of hea ic tem t proporti perature ona of t llyh to the e fuel throughout the domain. The mixture with EQ of 0.4 only reaches about twice the temperature within the throughout amount of f the dom uel introduced i ain. The mi nto the com xture with EQ of 0. bustion. Fi 4 onl gure y rea 16 show ches about twi s the static tem ce the tempera perature of t ture wi he fuel thin combustor. The temperature for the stoichiometric mixture, EQ 1.0, increases to approximately 3.3 times the combustor. The tempera throughout the domain. The mi ture f xture with EQ of 0. or the stoichiometr 4 onl ic mi y rea xture, EQ 1.0, ches about twi inccrea e the tempera ses to approxima ture withi teln y within the constant diameter section where it peaks at a ratio of 4.3 in the throat. Comparing the 3. the combustor. The tempera 3 times within the constant diam ture fet or the er secstoi tion where it chiometric mi peaks xture, EQ 1.0, at a ratio of 4i.n 3 crea in th ses to e throat approxima . Comparin telg y four-dierent fuel-air mixtures (Figure 17a–d), there is only temperature increase within the free-jet. t 3. h3 t e fou ime rs w -difife th rent in th fe const uel-air a mixt nt diam ures et (F er s igure ection where it 17a–d), there pe is aks only at a r temperature atio of 4.3 in t increase w he throat ithin . Com the free paring - The recirculation zone remains at a relatively low temperature of 1500 K compared to the 5000 K jet the fou . The rrecirc -different ulation fuel zone -air mixt remures ains ( at F ig a r ure elat 1i7 v ae –ly d low ), there tem is ponly erature temperature of 1500 K in co crease w mpared t ithin o the 5 the free 000 K - observed by the NASA study. The low temperatures at the nozzle part may be attributed to the observed by jet. The recirc the NA ulation SA zone st rem udy. a The ins at low a re temperatur latively lowe tse m at the nozzl perature of e p 15a 0rt m 0 K co ay mbe a pared ttri tb outed to the the 5000 K presence of reverse flow at the pressure outlet at the lower equivalence ratios of 0.4 and 0.6 at least, presence of r observed by ethe NA verse flow at the pressure SA study. The low outlet at temperatur the lower e es at the nozzl quivalee p nce rat art m ioa s of y be a 0.4 and ttributed to the 0.6 at least, but further work is clearly required in this area. but further presence of r we overse flow at the pressure rk is clearly required in this outlet at area. the lower equivalence ratios of 0.4 and 0.6 at least, but further With EQ 0.4, work is c most of the thermal learly required in this energ are y rele a. ase occurs in the divergent part of the combustion chamber. Th With EQ 0.4, e energ most of the thermal y release ends appr energ oximately y release halfway down occurs in the di the combustio vergent pan rt of the combustion chamber, meaning t chamber. Th hat all the feu energ el is cons y rele um ase end ed prior s appr to reach oximately ing the throa halfway down t. With the hi the combustio gh fueln -t chamber, m o-air mixtures, the eaning thermal ener that all the fugy r el ise cons lease o umccurs from ed prior to injection reaching unti the throa l the throa t. With the hi t of the combusti gh fuel-to-a on cha ir mixtures, the mber. The tthermal ener emperature igy r n the inl elease et ca occurs from n be observe injection d to be reunti lativl the throa ely low in al t of l four the combusti cases. on chamber. The temperature in the inlet can be observed to be relatively low in all four cases. Aerospace 2019, 6, 135 15 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 15 of 18 Aerospace 2020, 7, x FOR PEER REVIEW 15 of 18 Figure 16. Figure 16. Variation of t Variation of temperatur emperature with axia e with axiall distance distance.. Figure 16. Variation of temperature with axial distance. (a) (b) (a) (b) (c) (d) (c) (d) Figure 17. (a) Temperature contour for equivalence ratio of 0.4, (b) temperature contour for Figure 17. (a) Temperature contour for equivalence ratio of 0.4, (b) temperature contour for equivalence Figure 17. (a) Temperature contour for equivalence ratio of 0.4, (b) temperature contour for equivalence ratio of 0.6, (c) temperature contour for equivalence ratio of 0.8, (d) temperature contour ratio of 0.6, (c) temperature contour for equivalence ratio of 0.8, (d) temperature contour for equivalence equivalence ratio of 0.6, (c) temperature contour for equivalence ratio of 0.8, (d) temperature contour for equivalence ratio of 1.0. ratio of 1.0. for equivalence ratio of 1.0. 4.2.4. Pressure With EQ 0.4, most of the thermal energy release occurs in the divergent part of the combustion 4.2.4. Pressure Figures 18a–d and 19 plot the pressure drops throughout the combustion chamber. These results chamber. The energy release ends approximately halfway down the combustion chamber, meaning that Figures 18a–d and 19 plot the pressure drops throughout the combustion chamber. These results display the loss in performance of the engine. The pressure does show some fluctuation near the all the fuel is consumed prior to reaching the throat. With the high fuel-to-air mixtures, the thermal display the loss in performance of the engine. The pressure does show some fluctuation near the throat. The slight pressure increase in the throat followed by a pressure drop in the nozzle is expected energy release occurs from injection until the throat of the combustion chamber. The temperature in throat. The as the flow is slowed down slight pressure increase before re in the -acceler throat ating in follow the nozzle ed by a pressure . It is expected to have a pressur drop in the nozzle is ex e drop pected the inlet can be observed to be relatively low in all four cases. as ac the flow is slowed down ross the combustion chamber due to t before re-acc heler e foating in rmation th of shoc e nozzle k train . It is s within expected to have a pressur it. The presence of shoc e drop k formation can even promote mixing; however, the trade-off of efficient mixing comes with the across the combustion chamber due to the formation of shock trains within it. The presence of shock 4.2.4. Pressure decrease of total pressure recovery. formation can even promote mixing; however, the trade-off of efficient mixing comes with the Figures 18a–d and 19 plot the pressure drops throughout the combustion chamber. These results decrease of total pressure recovery. display the loss in performance of the engine. The pressure does show some fluctuation near the throat. The slight pressure increase in the throat followed by a pressure drop in the nozzle is expected as the flow is slowed down before re-accelerating in the nozzle. It is expected to have a pressure drop across the combustion chamber due to the formation of shock trains within it. The presence of shock formation can even promote mixing; however, the trade-o of ecient mixing comes with the decrease (a) (b) of total pressure recovery. (a) (b) (c) (d) (c) (d) Aerospace 2020, 7, x FOR PEER REVIEW 15 of 18 Figure 16. Variation of temperature with axial distance. (a) (b) (c) (d) Figure 17. (a) Temperature contour for equivalence ratio of 0.4, (b) temperature contour for equivalence ratio of 0.6, (c) temperature contour for equivalence ratio of 0.8, (d) temperature contour for equivalence ratio of 1.0. 4.2.4. Pressure Figures 18a–d and 19 plot the pressure drops throughout the combustion chamber. These results display the loss in performance of the engine. The pressure does show some fluctuation near the throat. The slight pressure increase in the throat followed by a pressure drop in the nozzle is expected as the flow is slowed down before re-accelerating in the nozzle. It is expected to have a pressure drop across the combustion chamber due to the formation of shock trains within it. The presence of shock Aerospace 2019, 6, 135 16 of 18 formation can even promote mixing; however, the trade-off of efficient mixing comes with the decrease of total pressure recovery. (a) (b) Aerospace 2020, 7, x FOR PEER REVIEW 16 of 18 (c) (d) Figure 18. (a) Pressure contour for equivalence ratio of 0.4, (b) pressure contour for equivalence ratio Figure 18. (a) Pressure contour for equivalence ratio of 0.4, (b) pressure contour for equivalence ratio of 0.6, (c) pressure contour for equivalence ratio of 0.8, (d) pressure contour for equivalence ratio of 1.0. of 0.6, (c) pressure contour for equivalence ratio of 0.8, (d) pressure contour for equivalence ratio of 1.0. Figure 19. Variation of pressure with axial distance. Figure 19. Variation of pressure with axial distance. 5. Conclusions 5. Conclusions The aim of this report was to adapt the Dual-Mode Free-Jet design, first tested by Trefny and The aim of this report was to adapt the Dual-Mode Free-Jet design, first tested by Trefny and Dippold III, and fit the FAME18 project propulsion requirements at cruise velocity Mach 8. The goal of Dippold III, and fit the FAME18 project propulsion requirements at cruise velocity Mach 8. The goal the FAME project was to design an advanced engine configuration for sustained commercial hypersonic of the FAME project was to design an advanced engine configuration for sustained commercial flight. The study was based on the prior work of the FAME team to provide improved designs of the hypersonic flight. The study was based on the prior work of the FAME team to provide improved inlet and nozzle for both the ramjet and scramjet, whilst also investigating the feasibility of the use of designs of the inlet and nozzle for both the ramjet and scramjet, whilst also investigating the a Dual-Mode Free-Jet combustion chamber design proposed by Trefny and Dippold. The dual-mode feasibility of the use of a Dual-Mode Free-Jet combustion chamber design proposed by Trefny and combustion engine or dual ramjet-scramjet is a system that operates as a ramjet at low supersonic Dippold. The dual-mode combustion engine or dual ramjet-scramjet is a system that operates as a speeds and then transforms to a scramjet at higher Mach numbers between 6 to 8. This design also ramjet at low supersonic speeds and then transforms to a scramjet at higher Mach numbers between benefits from being useable over a wide range of velocities while keeping reasonable performances. 6 to 8. This design also benefits from being useable over a wide range of velocities while keeping The 2010 numerical investigation by Trefny and Dippold III, was limited to hydrocarbon fuel reasonable performances. operation and challenges were encountered with extreme static temperatures within the combustion The 2010 numerical investigation by Trefny and Dippold III, was limited to hydrocarbon fuel chambers. This report adapted the Dual-Mode Free-Jet design to fit the FAME propulsion requirements operation and challenges were encountered with extreme static temperatures within the combustion at cruise velocity Mach 8 while using hydrogen instead of hydrocarbon-based fuels. chambers. This report adapted the Dual-Mode Free-Jet design to fit the FAME propulsion In this Computational Fluid Dynamics (CFD) investigation the k-" model was employed. The Y+ requirements at cruise velocity Mach 8 while using hydrogen instead of hydrocarbon-based fuels. value was maintained at less than 1. A Species Transport Laminar Finite Rate model was employed to In this Computational Fluid Dynamics (CFD) investigation the k-ε model was employed. The model the volumetric reaction combustion process. The chemical mechanisms were integrated into Y+ value was maintained at less than 1. A Species Transport Laminar Finite Rate model was the Fluent ANSYS software package using the CHEMKIN tool (used for solving complex chemical employed to model the volumetric reaction combustion process. The chemical mechanisms were kinetics problems). integrated into the Fluent ANSYS software package using the CHEMKIN tool (used for solving The Dual-Mode Free-Jet design was validated against numerical results provided by NASA at complex chemical kinetics problems). Mach 8. The model captured well all of the salient flow features such as flow deceleration (although The Dual-Mode Free-Jet design was validated against numerical results provided by NASA at not quite as dramatically as in the NASA model) and the stagnating recirculation zone separated by Mach 8. The model captured well all of the salient flow features such as flow deceleration (although not quite as dramatically as in the NASA model) and the stagnating recirculation zone separated by an intermediate zone of flow in the model, with the flow in both cases reattaching at the throat at a speed above Mach 1. In addition, the extreme temperatures in the recirculation region observed in the NASA numerical experiment were also successfully replicated in the validation study. The normal shock wave formation at the entrance of the inlet could not be prevented in this study. A way of preventing it would be implementing a variable inlet and nozzle height. Although, differences exist in the Mach and temperature plots, these are relatively minor, resulting in sufficient validations for further analysis. This new design allows the ramjet and scramjet engines to be combined into a single unit allowing for a volume reduction of 53% compared to prior work by the same group in the field (17). The aircraft is then capable to be run over a large range of Mach speeds using a single combustion chamber whilst allowing for significant space saving and reduced maintenance cost. As this study was the first to peer review the design, a series of challenges were faced when modelling the combustion chamber. Aerospace 2019, 6, 135 17 of 18 an intermediate zone of flow in the model, with the flow in both cases reattaching at the throat at a speed above Mach 1. In addition, the extreme temperatures in the recirculation region observed in the NASA numerical experiment were also successfully replicated in the validation study. The normal shock wave formation at the entrance of the inlet could not be prevented in this study. A way of preventing it would be implementing a variable inlet and nozzle height. Although, dierences exist in the Mach and temperature plots, these are relatively minor, resulting in sucient validations for further analysis. This new design allows the ramjet and scramjet engines to be combined into a single unit allowing for a volume reduction of 53% compared to prior work by the same group in the field (17). The aircraft is then capable to be run over a large range of Mach speeds using a single combustion chamber whilst allowing for significant space saving and reduced maintenance cost. As this study was the first to peer review the design, a series of challenges were faced when modelling the combustion chamber. Being the first study to study the unique combustion chamber design, this study provides useful and time saving insights for subsequent studies to develop upon the results. The size reduction of the combustion system for this aircraft as well as the sustainability improvement by switching from ethylene to hydrogen fuel are significant progress which encourages future advancement in research regarding this design. Author Contributions: M.C. was the research student that conducted the detailed study and wrote the paper assisted by A.A.S., who also edited the paper. A.P. conceived of the project, created the layout of the investigations, and checked the computational outcome of the resultant modelling eort and subsequent discussion. Funding: This research received no external funding. Conflicts of Interest: The authors declare no conflict of interest. References 1. Anderson, J.D. Hypersonic and High Temperature Gas Dynamics, 1st ed.; McGraw Hill College: New York, NY, USA, 1988. 2. Jonhson, R.P. An Emperical Method for the Prediction of Airplane Drag Divergence Mach Number; The RAND Corporation: Santa Monica, CA, USA, 1953. 3. Calvert, B. Flying Concorde: The Full Story, 3rd ed.; Airlife Publishing: Shrewsbury, UK, 2002. 4. Colonna, G.; Bonelli, F.; Pascazio, G. Impact of fundamental molecular kinetics on macroscopic properties of high-enthalpy flows: The case of hypersonic atmospheric entry. 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Baidya, R.; Pesyridis, A.; Cooper, M. Ramjet Nozzle Analysis for Transport Aircraft Configuration for Sustained Hypersonic Flight. Appl. Sci. 2018, 8, 574. [CrossRef] 18. Ridgway, A.; Sam, A.; Pesyridis, A. Modelling a hypersonic single expansion ramp nozzle of a hypersonic aircraft through parametric studies. Energies 2018, 11, 3449. [CrossRef] 19. Currant, A.T.; Stull, F.D. The Utilisation of Supersonic Combustion Ramjet Systems at Low Mach Numbers; RTD-TDR-63-4097; Research and Technology DIV Bolling AFB DC: Washington, DC, USA, 1964. 20. Smart, M.K. Comparison between Hydrogen and Methane Fuels in a 3-D Scramjet at Mach 8; The University of Queenland: Brisbane, Australia, 2016. 21. El-Sayed, A.F. Pulsejet, Ramjet and Scramjet Engine. In Fundamentals of Aircraft and Rocket Propulsion; Springer: London, UK, 2016; p. 382. 22. Solomon, S.; Rosenlof, K.H.; Portmann, R.W.; Daniel, J.S.; Davis, S.M.; Sanford, T.J.; Plattner, G. 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In Proceedings of the 17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference, San Francisco, CA, USA, 11–14 April 2011. 26. Choi, J.J.; Ghodke, C.D.; Menon, S. Large Eddy Simulation of Cavity Flame Holding in a Mach 2.5 cross Flow. In Proceedings of the 48th AIAA Aerospace Sciences Meeting including the new Horizons Forum and Aerospace Exposition, Orlando, FL, USA, 4–7 January 2010. © 2019 by the authors. Licensee MDPI, Basel, Switzerland. This article is an open access article distributed under the terms and conditions of the Creative Commons Attribution (CC BY) license (http://creativecommons.org/licenses/by/4.0/).
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