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Home , Hypersonic speed, Jet fuel, Mach number, Pulsejet, Ramjet, Scramjet

Article Modelling of a Dual--Mode Free-Jet System

, Maxim Cooper †, Ashish Alex Sam † and Apostolos Pesyridis * † College of and Design, Brunel University , Uxbridge, London UB8 3PH, UK; [email protected] (M.C.); [email protected] (A.A.S.) * Correspondence: [email protected]; Tel.: +44-18-9526-7901 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 design, first tested in 2010 by NASA, is being adapted to run on fuel instead of ethylene while addressing the excessive thermal heat load. This study is part of the FAME ( at Mach Eight) project, with the primary objective to design and analyse the configuration for a hypersonic commercial . 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 different components that make-up the aircraft and how their designs affect one another, from the generation component, fuel, propulsion system and control surfaces [1]. The invention of safe and reliable jet made any part of the world accessible within 24 h. The and the Tu-144 have been the only 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 , Mach 0.8 to 0.9, is considered inefficient, due to the sudden increase of when travelling above the 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 different engine such has , and . As these types of propulsion technologies are fundamentally different to the standard and 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 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 going faster has brought new engineering challenges. The accepted convention in that defines hypersonic flow is where the flow speeds are greater than Mach 5. At these speeds, the physical behavior of air is very different to what is encountered at subsonic. When an object is flying through at the atmosphere, at supersonic or , 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 increase, and . The viscosity of the fluid plays an intricate role with the of the fluid. The high 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-off (with ) and during supersonic cruise were 170 kN and 45 kN, respectively. The engines were run on A1 , 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 systems. At cruise speeds, the 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 . To improve the efficiency 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 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-off 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-off to cruise speeds and back to landing. However, the engine emitting large amount of noise and pollutants due to its poor efficiency 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 from the compressor straight into the afterburner, which makes the engine behave like a . Effectively the afterburner would become the combustion chamber. The effect was to improve the efficiency 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 off 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 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 (H2O), 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 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 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 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 fuel rocket technology [8]. The aircraft must carry the oxidizer as well, since the amount of in the atmosphere at high altitudes is insufficient 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 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 , Rolls-Royce and called E-Fan X Hybrid promises greener travel [9]. British , Limited is working on a precooled 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 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 , 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 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 . The study will also investigate how effectively altering the fuel to air ratio addresses the excessive temperatures within Aerospace 20202019,, 76,, x 135 FOR PEER REVIEW 4 4of of 18 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].

2.2. Dual-Mode Dual-Mode Combustion Combustion Engines Engines AsAs is is true true for the body of the aircraft, differen differentt engine technologies func functiontion best best at at different different altitudesaltitudes and and velocities. None None of of the the current current commercial commercial aircraft aircraft are are fitted fitted with engines capable of propelling them to velocities beyond Mach 0.85. This This second second challenge challenge wi willll be be addressed by by using differentdifferent engine technologies, ramjet and scramjet. Ho However,wever, for for the the aircraft aircraft to to be be a a viable viable option option for for companies, it must take off like any other aircraft from a conventional runway. This is airline companies, it must take off 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 are limited to by the turbine drives the compressor. Where turbofans are limited to Mach 0.85, , 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-off 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. Figure1 describes the typical flight profile of the plane from take-o ff 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-off 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 different 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.In this design, The engine a dual-mode configuration combustion includes, chamber two turbojet which enginescan be used with as two both ramjet ramjet inlets and to scramjet channel isthe employed. air into the The turbojet engine engines. configuration When, the includes, aircraft hastwo reached turbojet a velocityengines ofwith Mach two 3 and ramjet an altitude inlets to of channel10,000 m, the part airof into the the air turbojet is channeled engines. into When, the dual-mode the aircraft combustion has reached chamber a velocity which of isMach configured 3 and an in altitude of 10,000 m, part of the air is channeled into the dual-mode combustion chamber which is

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configured in ramjet operation. The turbojets and ramjet work hand in hand until Mach 4 is reached ramjetat an altitude operation. of The20,000 turbojets m. At this and ramjetstage the work turboj handet inceases hand to until work. Mach All 4the is reachedair entering at an the altitude inlets of is 20,000routed m. into At the this ramjet. stage Finally, the turbojet once ceasesa speed to of work. Mach 6 All is reached, the air entering the dual-mode the inlets combustion is routed chamber into the ramjet.is switched Finally, from once ramjet a speed to scramjet of Mach mode. 6 is reached, Scramjet the mode dual-mode will remain combustion in operation chamber for the is switchedmost part fromfor the ramjet flight to where scramjet it will mode. reach Scramjet its maximum mode will velocity remain of inMach operation 8 at altitude for the 30,000 most partm. for the flight whereAnother it will reach improvement its maximum to the velocity body from of Mach the 8previous at altitude work 30,000 [13] m. is the use of a series of “flaps” or “doors”Another used improvement to channel tothe the airflow body into from the the diff previouserent engines. work [ 13These] is theare useshown of a as series dotted of “flaps”lines in orFigure “doors” 2. These used todevices channel prevent the airflow unnecessary into the drag different formation engines. by Theseletting are air shown enter engines as dotted which lines are in Figurenot in2 use.. These For devicesexample, prevent at Mach unnecessary 8, the ramjet drag an formationd turbojet byinlets letting are airnot enter being engines used and which thus are can not be in“closed use. For off” example, to make at Machthe body 8, the of ramjet the aircraft and turbojet more inletsaerodynamically are not being streamlined. used and thus The can need be “closed of two oturbojetff” to make engines the bodyis required of the for aircraft safety more reasons. aerodynamically In the situation streamlined. where one turbojet The need engine of two fails turbojet at low enginesaltitude, is i.e., required take-off for and safety landing reasons. phase, In thea second situation engine where is required one turbojet to ensure engine the fails safe at landing low altitude, of the i.e.,aircraft. take-o ff 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 subsonicThe or design supersonic of the combustion.combustion Thischamber characteristic allows for makes the combustion it suitable for chamber a wide to range be used of operating at either velocities.subsonic or The supersonic dual-mode combustion. supersonic This combustion characteristic engine makes design it issuitable patented for bya wide Currant range and of Stulloperating [20]. Duringvelocities. the The use dual-mode of ramjet, the supersonic airflow needs combustion two throats; engine the design first throatis patented is to reduceby Currant the flow and Stull speed [20]. to subsonicDuring the level use for of combustion, ramjet, the andairflow the secondneeds two one throat is for supersonics; the first throat acceleration is to reduce in the nozzlethe flow (Figure speed3 ).to However,subsonic level for scramjet for combustion, the throats and are the not second needed one asis itfor can supersonic be seen inacceleration Figure4. in the nozzle (Figure 3). However, for scramjet the throats are not needed as it can be seen in Figure 4.

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Figure 3. Free-jet dual-mode 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 difference between a ramjet and scramjet engine is their operation speed range and their combustionThe difference process velocities.betweenFigure a4. Theramjet Free-jet operation and dual-mode scramjet speed combustor range engine of isin aramjet supersonictheir operation lies betweenmode. speed Mach range 2 and and 5 whiletheir thecombustion combustion process process velocities. occurs The at subsonic operation velocity. speed range A scramjet of a ramjet can functionlies between from Mach Mach 2 and 4 upwards 5 while whilethe combustionThe the combustiondifference process between process occurs a occurs atramjet subsonic at and lower scramjetvelocity. supersonic enA ginescramjet velocity. is their can operationfunction fromspeed Mach range 4 upwardsand their whilecombustionThere the combustion are process technical velocities. process challenges occurs The with operation at this lower type speed supersonic of hybrid range engine velocity.of a ramjet configuration, lies between due Mach to the 2 modulationand 5 while ofthe the combustionThere thermal are location, technicalprocess fueloccurschallenges distribution, at subsonic with flame thisvelocity. type holding Aof scramjet hybrid and the caenginen actual function configuration, ignition from of Mach the due fuel4 upwards to in the combustionmodulationwhile the combustion of chamber the thermal inprocess such location, occurs a large fuel at cross-section distribution,lower supersonic [flame13]. velocity. Whenholding used and atthe high actual operating ignition velocities,of the fuel Machin theThere 5combustion and aboveare technical ,chamber the supersonic challenges in such combustion a large with cross-sect this occurs typeion in of an[13]. hybrid unconfined When engine used supersonic at configuration, high operating free-jet surroundeddue velocities, to the byMachmodulation a subsonic 5 and of combustion. above,the thermal the location, Thesupersonic free-jet fuel iscombustion distribution, channeled towards occursflame holdingin a variablean andunconfined throatthe actual area supersonic ignition nozzle. of As thefree-jet see fuel in Figuresurroundedin the 4combustion, the supersonicby a subsonic chamber propulsive combustion. in such stream a large The is cross-sectfree-jet not in is contact channeledion [13]. with When towards the wallused a of atvariable thehigh combustion operating throat area velocities, chamber. nozzle. ToAsMach adjustsee in5 Figure theand nozzle above, 4, the throat supersonicthe crosssupersonic section propulsive area,combustion stream a control isoccurs systemnot in contactin is neededan withunconfined to th managee wall supersonic of the the reattachment combustion free-jet ofchamber.surrounded the flow To when byadjust a subsonic in scramjetthe nozzle combustion. operation. throat cross The Thus, free-jetsectio then useis area, channeled of a dual-modecontrol towards system engine a variable is needed could throat beto beneficialmanage area nozzle. the as itreattachmentAs off seeers in up Figurea wayof the4,of the flow building supersonic when up in to propulsivescramjet the required operat stream Machion. is Thus, not 8 speed in the contact valueuse of with through a dual-mode the wall an integration of engine the combustion could of two be enginebeneficialchamber. systems. Toas itadjust offers This the leadsup nozzle a toway space, throat of weightbuilding cross andsectio up maintenance ton area,the required a control cost Mach savings. system 8 speedis Having needed value less to combustionthroughmanage thean chambersintegrationreattachment increases of twoof the engine theflow space systems.when for inextra scramjetThis leads payload operat to space, orion. fuel weightThus, and the reduces and use maintenance of the a dual-mode needto cost maintain savings. engine multiplecould Having be combustionlessbeneficial combustion as chambers. it offerschambers up The a localizedincreases way of building heatingthe space e ffup ectsfor to extra arethe negated requiredpayload due Machor to fuel the 8 and flowspeed reduces not value being thethrough in need contact anto withmaintainintegration the combustion multiple of two enginecombustion chamber systems. walls. chambers. This A secondary leads The to localized space, flow weight of heating air can and beeffects maintenance introduced are negated into cost the savings.due engine to the Having to flow act asnotless a being coolingcombustion in agentcontact chambers whilst with increasingthe increases combustion the the pressure. chamberspace for However, wa extralls. Apayload thissecondary consideration or fuelflow and of willair reduces can not be be theintroduced studied need into thisintomaintain paper.the engine multiple to act combustion as a cooling chambers. agent whilst The increasinglocalized heating the pressure. effects However,are negated this due consideration to the flow willnot beingnotThe be high-speed in studied contact in airwith this enters paper.the combustion the engine through chamber the wa inletlls. A where secondary it is compressed flow of air through can be introduced a series of shockwaves.into Thethe enginehigh-speed As itto is act compressed, air as entersa cooling the it isagentengine slowed whilst through down increasing to the subsonic inlet the where speeds pressure. it inis thecompressed However, case of a ramjetthis through consideration engine a series and remainsofwill shockwaves. not supersonicbe studied As ininit scramjetisthis compressed, paper. engine. it However, is slowed the down flow isto still subsonic slowed speeds down to in improve the case the of eaffi ramjetciency ofengine theThe combustion and high-speed remains chamber. supersonic air enters Once inthe scramjet in engine the combustion engine.through Ho the chamber,wever, inlet wherethe the flow fuel it isis is stillcompressed injected slowed and down through mixed to improve with a series the air.theof shockwaves. Inefficiency the case of of theAs ramjet, itcombustion is flamecompressed, holders chamber. it (igniters) is Onceslowed mayin downthe be combustion required to subsonic to chamber, maintain speeds athin steadye thefuel case iscombustion injected of a ramjet and of themixedengine fuel. with and In scramjettheremains air. In engine,supersonic the case the of fuelin ramjet, scramjet automatically flame engine. hold ignites ersHo wever,(igniters) due the to may the flow extreme be is required still slowed temperatures to maintain down ofto theaimprove steady flow enteringcombustionthe efficiency the combustionof of the the fuel.combustion chamber.In scramjet chamber. Finally, engine, for Once both th ine engines, thefuel combustion automatically the flow chamber, is expandedignites th duee andfuel to ejectedis theinjected extreme though and thetemperaturesmixed nozzle. with Thethe of air. nozzle’sthe In flow the role caseentering isof toramjet, convertthe combustion flame the hold pressureers chamber. (igniters) and thermal Finally, may be energy for required both of theengines, to maintain air and the fuel aflow steady into is kineticcombustion energy of which the fuel. in turn In generatesscramjet thrust.engine, the fuel automatically ignites due to the extreme temperatures of the flow entering the combustion chamber. Finally, for both engines, the flow is

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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 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 Triffny 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 Table1, below.

Table 1. Comparison of fuel properties.

Fuel Energy/Mass Energy/Volume Hydrogen 116.7 8.2 71 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 offers. Its only by-product is water vapor (H2O). 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-off to cruise velocity, therefore relieving the need for for the other engines. The water vapour produced by the combustion of hydrogen does have a polluting effect at very high altitude [22]. The presence of water vapor in the stratosphere, has a warming effect of the troposphere, hence being a contributor in global warming [23].

2.4. Injection and Combustion The challenge with supersonic combustion is achieving sufficient 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 effect, heat addition from adding fuel into the supersonic flow may result in a decrease of the . An excessive decrease in pressure will impact the efficiency 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.

3.3. ComputationalComputational MethodologyMethodology TheThe computationalcomputational simulationssimulations havehave thethe goalgoal toto simulatesimulate andand quantifyquantify thethe performancesperformances ofof thethe adaptedadapted Dual-ModeDual-Mode Free-Jet Free-Jet design design to tofit fitthe the propulsionpropulsionrequirements requirements at at cruise cruise velocity velocity Mach Mach 8 8 whilewhile usingusing hydrogenhydrogen insteadinstead ofof hydrocarbon-basedhydrocarbon-based fuels. fuels. CommercialCommercialsoftware softwareAnsys AnsysFluent Fluentwas wasused usedto to performperform thethe simulations.simulations. InvestigationInvestigation onon thethe eeffectffect adjustingadjusting thethe fuelfuel toto airair ratio,ratio, fromfrom equivalenceequivalence ratioratio 0.40.4 toto 1.0,1.0, hashas onon the the temperatures temperatures within within the the combustion combustion chamber chamber will will also also be be studied. studied.

3.1.3.1. GeometryGeometry TheThe geometrygeometry of of thethe combustor,combustor, used used for for all all cases, cases, was was identical identical to to the theone oneused usedin in thethe 20102010 studystudy byby TrefnyTrefny andand DippoldDippold IIIIII [13[13]] (Figure(Figure5 5and and Table Table2). 2). The The dimensions dimensions were were kept kept identical identical for for the the validationvalidation case case and and were were modified modified to to suit suit the the FAME FAME objectives. objectives. The The dotted dotted line line between between A and A and K is K the is symmetrythe symmetry line. line. The The step step at D at and D and C isC tois to induce induce flow flow separation separation from from the the wall. wall. The The curve curve through through pointpoint G–H–IG–H–I isis defineddefined withwith twotwo tangenttangent constraintsconstraints toto thethe neighboringneighboring lineline F–GF–G andand I–JI–J andand fixedfixed heightheight forfor the throat, point point H. H. The The nozzle nozzle section section is isnot not part part of the of the combustor combustor final final design, design, but its but only its onlypurpose purpose is to isdemonstrate to demonstrate expansion expansion of the of flow the flow after after passing passing the throat the throat (H) during (H) during simulation. simulation. The Theramjet ramjet andand scramjet scramjet nozzles designed designed areare dimensioned dimensioned to tobe befitted fitted at atpoint point H. H. The The injection injection system system is ispositioned positioned between between points points A A an andd B, B, as as shown shown in in Figure Figure 6.6 .

FigureFigure 5.5. CombustionCombustion chamberchamber 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.

CombustionCombustion Chamber Chamber Dimensions Dimensions Injector Injector Dimensions Dimensions PointPoint X (mm) X (mm) Y (mm) Y (mm) Point Point X(mm) X (mm) Y(mm) Y (mm) AA 0.00.0 0.00.0 LL 0.0 0.017.20 17.20 BB 0.00.0 119.5119.5 MM 0.0 0.022.60 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 ED 595.2111.0 351.0131.4 PO 0.0 0.062.40 96.80 FE 1297.1595.2 351.0351.0 QP 0.0 0.096.80 102.20 GF 1407.51297.1 248.2 351.0 Q 0.0 102.20 HG 1533.11407.5 196.7 248.2 I 1622.6 216.0 JH 2200.01533.1 500.0 196.7 KI 2200.01622.6 0.0 216.0 J 2200.0 500.0 K 2200.0 0.0

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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 dimensiondimension of of the the injectors injectors had tohad be drasticallyto be drastically reduced reduced in size to accommodatein size to accommodate the replacement the replacement fuel, from ethylene to hydrogen due to the higher calorific content. In a ramjet-scramjet replacementfuel, from ethylene fuel, from to ethylene hydrogen to due hydrogen to the due higher to the calorific higher content. calorific Incontent. 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 combustion the combustionchamber chamber can be asymmetrically can be asymmetrically shaped, meaning shaped, itmeaning does not it needdoes tonot be need cylindrical to be cylindrical or axially symmetrical [22]. Figure 7 shows the asymmetrical design of the Dual-Mode cylindricalor axially symmetricalor axially symmetrical [22]. Figure [22].7 shows Figure the 7 shows asymmetrical the asymmetrical design of design the Dual-Mode of the Dual-Mode Free-Jet Free-Jet combustion chamber. Free-Jetcombustion combustion chamber. 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 Figure 7. 3D view of dual-mode rectangular combustion chamber. 3.2. Boundary Types 3.2. BoundaryThe inlet Types (fuel and air) and outlet boundaries were specified using pressure. The computational domainThe was inlet halved (fuel and employing air) and outlet symmetry boundaries conditions were specified so as to using reduce pressure. the computational The computational effort. The inlet (fuel and air) and outlet boundaries were specified using pressure. The computational Thedomain boundary was halved conditions employing are described symmetry in Table conditions3. 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 PressureTable 3. (Pa) Boundary conditions Temperature. (K) Species Total Pressure (Pa) Temperature (K) Species Inlet Total Pressure 150,105 (Pa) Temperature 3012 (K) Species O2—21.15% InjectorInlet 150,105150,105 3012547 O2 C—21.15%2H4—100% OutletInlet 150,105 1580 3012 227 O2—21.15% 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 3.3. NumericalThere is no Model single turbulence model able to universally predict every turbulent flow. Choosing the correctThere model is canno single have significant turbulence impact model on able the accuracyto universally of the predict flow modelling. every turbulent The calculation flow. Choosing carried There is no single turbulence model able to universally predict every turbulent flow. Choosing outthe bycorrect NASA model used can a time-accurate, have significant fully impact implicit on codethe accuracy developed of the in-house flow modelling. [13]. In this The study calculation the k-ε the correct model can have significant impact on the accuracy of the flow modelling. The calculation modelcarried was out employed.by NASA used RANS a time-accurate, models are applicable fully implicit to a code wide developed range of applications in-house [13]. while In this proving study carried out by NASA used a time-accurate, fully implicit code developed in-house [13]. In this study appropriatelythe k-ε model accurate was employed. for this study. RANS Density models based are models applicab arele recommended to a wide range for supersonicof applications simulation 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

AerospaceAerospace 202020192020,,, 767,,, x 135x FORFOR PEERPEER REVIEWREVIEW 101010 ofof of 1818 supersonicsupersonic simulationsimulation sincesince theythey provideprovide goodgood convergenceconvergence propertiesproperties asas opposedopposed toto pressure-basedpressure-based models.models.since they ForFor provide thethe solutionsolution good convergence ofof thethe governinggoverning properties equatiequati as opposedons,ons, anan to pressure-basedimplicitimplicit approachapproach models. withwithFor thethe the Roe-fluxRoe-fluxsolution differencedifferenceof the governing splittingsplitting equations, forfor calculatingcalculating an implicit thethe convectiveconvective approach with fluxflux thewaswas Roe-flux used.used. TheThe di ff flow,flow,erence turbulentturbulent splitting kinetickinetic for calculating energy,energy, andandthe convective turbulentturbulent dissipationdissipation flux was used. werewere The setset flow, asas secondsecond turbulent orderorder kinetic upupwindwind energy, toto obtainobtain and turbulent moremore accuaccu dissipationraterate solutions.solutions. were SingleSingle set as step,step,second volumetricvolumetric order upwind reactionsreactions to obtain andand morefinitefinite accurateraterate reactireacti solutions.onon forfor thethe Single turbulence-chemistryturbulence-chemistry step, volumetric reactions interactioninteractionand were finitewere enabledenabledrate reaction withwith forthethe the hydrogen–airhydrogen–air turbulence-chemistry mixturemixture (H(H interaction22 ++ OO22),), toto modelmodel were enabledthethe scramjetscramjet with combustion.combustion. the hydrogen–air TheThe Y+Y+ mixture valuevalue was(Hwas2 +maintainedmaintainedO2), to model lessless thethanthan scramjet 1.1. AA SpeciesSpecies combustion. TransportTransport The LaminarLaminar Y+ value FiniteFinite was maintainedRaRatete modelmodel less willwill than bebe employedemployed 1. A Species toto modelTransportmodel thethe Laminarvolumetricvolumetric Finite reactionreaction Rate model combustioncombustion will be processprocess employed wherewhere to model thethe turbulentturbulent the volumetric fluctuationsfluctuations reaction areare combustion ignored.ignored. ItIt usesprocessuses thethe where ArrheniusArrhenius the turbulent RateRate ofof ReactionReaction fluctuations Law.Law. are TheThe ignored. chemicalchemical It uses mechanismsmechanisms the Arrhenius areare integratedintegrated Rate of Reaction intointo FluentFluent Law. ANSYSTheANSYS chemical usingusingmechanisms thethe CHEMKINCHEMKIN are integrated tool.tool. ThisThis into softwaresoftware Fluent is ANSYSis usedused usingforfor solvingsolving the CHEMKIN complexcomplex tool.chemicalchemicalThis softwarekineticskinetics problems.problems.is used for solving complex chemical kinetics problems.

3.4.3.4. MeshMesh Mesh TheThe computational computational domaindomain waswas meshedmeshed usingusing aa quadrilateral quadrilateral meshmesh andand using using edge edge sizing sizing whereverwherever possible. possible. FigureFigure 88 showsshowsshows thethethe gridgridgrid patternpatternpattern generatedgeneratedgenerated tototo thethethe NASA NASANASA specifications. specifications.specifications. TheTheThe gridgridgrid shownshown inin the the figure figurefigure is is with with reducedreduced meshmesh densitydensity forfor betterbetter clarity.clarity. TheThe nodenode numbernumber usedused forfor validationvalidation was was in in accordance accordance with with numb numbernumberer used used by by NASA, NASA, approximatelyaapproximatelypproximately 392,000.392,000.

FigureFigure 8. 8. ComputationalComputational mesh.mesh.

TheThe gridgrid independenceindependence studystudy waswas carriedcarried carried outout out usingusing using ANSYSANSYS ANSYS Workbench,Workbench, Workbench, wherewhere 55 differentdifferent different meshmesh densitiesdensities werewere created, created, a a 50,000, 50,000, 70,000, 70,000, 110,000, 110,000, 177,000 177,000 and and 267,000. 267,000. TheThe 110-k110-k cellcell meshmesh showedshowed thethe fastestfastest convergenceconvergence toto steadysteady massmass flowflow flow rate rate of of 0 0 kg/s kgkg/s/s (Figure(Figure 9 9).9).). CoarserCoarserCoarser meshesmeshesmeshes providedprovidedprovided aaa veryveryvery unstableunstable valuevalue atat thethe beginningbeginning ofof thethe simulationsimulation andand wouldwould endend upup takingtaking aa greatergreater amountamount ofof timetime toto converge. converge. DenserDenser Denser meshesmeshes meshes providedprovided provided moremore more stablestable stable results;results; results; however,however, thethe convergence convergence would would requirerequire aa greater greater amountamount ofofof timetimetime to toto converge, converge,converge, particularly particularlyparticularly for forfor 267,000 267,000267,000 cells cellscells and andand above. ababove.ove. It was ItIt waswas decided decideddecided upon uponupon the thetheresults resultsresults of the ofof the meshthe meshmesh refinement refinementrefinement study studystudy to carry toto carrycarry all simulations alalll simulationssimulations with withwith a mesh aa meshmesh count countcount of 110,000 ofof 110,000110,000 cells. cells.cells.

FigureFigure 9.9. MeshMeshMesh refinementrefinement refinement study.study.

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3.5. Species Model The three-step reaction with 7 species for ethylene–air is described in Table4[ 23] and the one step reaction for hydrogen–oxygen is described in Table5.

Table 4. Ethylene–air chemical reaction and products.

Reactions Products Species Stoich Coefficient Rate Exponent Species Stoich Coefficient Rate Exponent

C2H4 1 1 CO 2 0 O2 1 1 H2 2 0 Pre exponential factor Activation energy Temperature exponent 2.10 1011 1.50 108 0 × × Species Stoich Coefficient Rate Exponent Species Stoich Coefficient Rate Exponent CO 2 1 CO 2 0 O2 1 1 Pre exponential factor Activation energy Temperature exponent 3.48 105 8.43 107 2 × × Species Stoich Coefficient Rate Exponent Species Stoich Coefficient Rate Exponent

H2 2 1 H2O 2 0 O2 1 1 Pre exponential factor Activation energy Temperature exponent 3.00 1014 0 1 × −

Table 5. Hydrogen–oxygen chemical reaction and products.

Reactions Products Species Stoich Coefficient Rate Exponent Species Stoich Coefficient Rate Exponent

H2 1 1 H2O 1 0 O2 0.5 1 Pre exponential factor Activation energy Temperature exponent 9.87 108 3.1 107 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 difference 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)( b)

FigureFigure 10.10. ComparisonComparison ofof Mach Mach contour; contour; (a ( )a computational) computational result, result, (b) ( experimentalb) experimental result result [17]. [[17]. 17].

TheThe extreme extreme temperatures in in the the recirculation recirculation regi regionregionon of of5000 5000 kelvin kelvin (9500 (9500 Rankin) Rankin) were were observed observed inin thethe the NASA NASANASA numericalnumericalnumerical experiment,experiment, Figure FigureFigure 11b. 1111b.b. This This was was successfully successfully replicated replicated in the in thevalidation validationvalidation study.study. TheThe The subsonic subsonicsubsonic recirculating recirculating flow flow flow can can can be be be seen seen seen to to beto be inbe in thein the the region region region of of 5000 of5000 5000 kelvin kelvin kelvin degrees. degrees. degrees. These These These high temperaturehighhigh temperaturetemperature causes causes ionizing ionizingionizing of atoms, ofof which atoms,atoms, iswhich undesirablewhich is isundesirable undesirable for combustion. for for combustion. combustion. This can beThis addressedThis can canbe bybe adjustingaddressedaddressed the byby amount adjustingadjusting of the fuel amountamount being introduced of of fuel fuel being being into in introduced thetroduced combustion into into the the chamber. combustion combustion Although chamber. chamber. di ffAlthougherences Although exist indifferencesdifferences the Mach exist andexist temperature in the Mach plots, andand the temperaturetemperature setup provides pl plots,ots, results,the the setup setup which provides provides provide results, suresults,fficient which which validations provide provide for furthersufficientsufficient analysis. validationsvalidations for furtherfurther analysis. analysis.

(a) (b) (a) (b) Figure 11. Comparison of temperature distribution, (a) computational result, (b) historical numerical Figureexperimental 11. Comparison result [17]. of temperature distribution, ( a) computational result, (b) historical numerical experimental result [[17].17]. 4.2. Effect of Equivalence Ratio on the Flow Characteristics 4.2. Effect Effect of Equivalence RatioRatio on the Flow Characteristics Following the validation run of the NASA design, the model was modified to meet the specificationsFollowing for thethe the validation validation present study. run run of theThe of NASA numericalthe NASA design, methodology design, the model the parameters was model modified was used tomodified meetare identical the specificationsto tomeet the the forspecificationsvalidation the present case, for study. except the present The that numerical the study. species The methodology and numerical materials parameters methodologyparameters used have parameters are been identical changed used to to theare reflect validationidentical the use to case, the exceptvalidationof hydrogen that case, the fuel. species except The effect andthat materialsthe of equivalence species parameters and ratio material (E haveQ)s onparameters been the changedflow characteristics have to reflectbeen changed the has use been ofto hydrogenreflectstudied the in fuel.use Theofdetail. hydrogen effect of equivalencefuel. The effect ratio of (EQ)equivalence on the flow ratio characteristics (EQ) on the flow has characteristics been studied in has detail. been studied in detail. 4.2.1.4.2.1. MachMach NumberNumber 4.2.1.As AsMach seenseen Number inin Figure 12a–d,12a–d, the the Mach Mach velocity velocity at atthe the inlet inlet is below is below Mach Mach 1.0. 1.0.This Thissuggests suggests that a that anormal normal shockshock perpendicular perpendicular to the to the flow flow has hasformed formed at the at very the veryentrance entrance of the ofinlet, the which inlet, causes which the causes As seen in Figure 12a–d, the Mach velocity at the inlet is below Mach 1.0. This suggests that a theflow flow to toslow slow down down as asexplained explained above. above. This This shock shock formation formation causes causes the theflow flow to go to into go into subsonic subsonic normal shock perpendicular to the flow has formed at the very entrance of the inlet, which causes the velocitiesvelocities inin thethe combustioncombustion chamber. chamber. In In essence, essence, the the scramjet scramjet engine engine is behaving is behaving as a as ramjet a ramjet where where flow to slow down as explained above. This shock formation causes the flow to go into subsonic combustioncombustion occurs at at subsonic subsonic flow flows.s. The The free-jet free-jet is at issubsonic at subsonic velocity velocity instead insteadof supersonic. of supersonic. The velocities in the combustion chamber. In essence, the scramjet engine is behaving as a ramjet where Thefree-jet free-jet is surroundedis surrounded by a by lower a lower speed speed static static regi regionon of flow. of flow. The Thefuel fuelcan canbe seen be seen to be to traveling be traveling combustion occurs at subsonic flows. The free-jet is at subsonic velocity instead of supersonic. The throughthrough thethe free-jetfree-jet atat a lower speed speed and and the the surroun surroundingding air air from from the the inlet. inlet. Looking Looking at the at theMach Mach free-jet is surrounded by a lower speed static region of flow. The fuel can be seen to be traveling ratioratio along along thethe axialaxial centerline distance distance of of the the combustor combustor (Figure (Figure 13), 13 significant), significant fluctuations fluctuations occur occur at at through the free-jet at a lower speed and the surrounding air from the inlet. Looking at the Mach EquivalenceEquivalence ratioratio 0.40.4 where little fuel fuel is is injected. injected. The The peak peak velocity velocity occurs occurs in the inthe divergent divergent part part of the of the ratiocombustor, along the between axial centerline 0.5 m and distance 2.2 m in ofall the cases. combustor However, (Figure only for 13), equivalence significant ratiofluctuations 0.4 does occur the at combustor, between 0.5 m and 2.2 m in all cases. However, only for equivalence ratio 0.4 does the EquivalenceMach reaches ratio 1.4 times0.4 where the inlet little speed. fuel is injected. The peak velocity occurs in the divergent part of the 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.

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(a) (b)

(c) (d)

Figure 12.12. ((aa)) Mach Mach number number contour contour for for equivalence equivalence ratio ratio of of 0.4, ( b) Mach number contour for equivalence ratioratio of of 0.6, 0.6, (c )( Machc) Mach number number contour contour for equivalence for equivalence ratio ofratio 0.8, of (d )0.8, Mach (d) number Mach number contour forcontour equivalence for equivalence ratio of 1.0. ratio of 1.0.

Figure 13. Variation of Mach ratio with axial distance.

The equivalence ratio ratio mixture mixture 0.6 0.6 peeks peeks at at a arati ratioo 1, 1,meaning meaning its its maximum maximum speed speed is equal is equal to its to itsinlet inlet speed speed (Mi). (Mi). EQ 0.8 EQ and 0.8 1.0 and show 1.0 poor show Mach poor ratios Mach below ratios 1 belowthroughout 1 throughout the combustion the combustion chamber. chamber.The Mach Theplots Mach in Figureplots 12a–d in Figure shows 12 a–dthe flow shows velo thecities flow to velocities be above toMach be above1 prior Mach to reaching 1 prior the to reachingthroat, hence the throat, allowing hence the allowing flow to theaccelerate flow to in accelerate the nozzle. in the However, nozzle. However,the expected the expectedMach velocity Mach velocitybetween betweenMach 2.5 Machto 3.0 required 2.5 to 3.0 for required the scramjet for the nozzle scramjet to produce nozzle enough to produce thrust enough was not thrust achieved. was notAs the achieved. amount As of the fuel amount is increased, of fuel isfrom increased, equivalence from equivalence ratio 0.4 to ratio 1.0, 0.4the toMach 1.0, thevalue Mach remains value remainsunchanged. unchanged. This suggestsThis suggests that a leanthat fuel-air a lean fuel-airmixture mixture would wouldhave an have as angood as goodeffect eofffect thrust of thrust as a asstoichiometric a stoichiometric mixture. mixture.

4.2.2. Velocity Similar to to the the Mach Mach plots, plots, the the velocity velocity plots plots in Fi ingure Figure 14a–d 14a–d show show a relative a relativelyly fast subsonic, fast subsonic, free- free-jetjet flow flowsurroundedsurrounded by a byslower, a slower, recirculating recirculating zone zoneof airof above air above it. The it. freestream The freestream jet reattaches jet reattaches at the atthroat the throatwhere wherein reaches in reaches a velocity a velocity of about of about 650 to 650 850 to m/s 850 depending m/s depending on the on fuel the fuelcontent content of the of themixture. mixture. The relationship The relationship between between fuel-to-air fuel-to-air ratio and ratio flow and accelera flow accelerationtion becomes becomes clear in clear the nozzle in the nozzleas shown as shownby Figure by Figure15. With 15. an With EQ an of EQ 0.4 ofand 0.4 0.6, and the 0.6, lean the leanmix mixof fuel of fueldoes does not notallow allow the theair airto toaccelerate accelerate above above 960 960 m/s m in/s the in the nozzle. nozzle. However, However, with with higher higher fuel fuel mixtures, mixtures, EQ EQ1.0 and 1.0 andEQ 0.8, EQ 0.8,the theflow flow acceleratesaccelerates beyond beyond 1000 1000 m/s. m Figure/s. Figure 14b,c 14 showsb,c shows twotwo dark dark red redstreaks streaks of fast of fastmoving moving flow flow in the in thenozzle nozzle at EQ at EQ0.8 0.8and and EQ EQ 1.0, 1.0, which which are are not not present present at atEQ EQ 0.4 0.4 and and are are only only faintly faintly visible visible at at EQ EQ 0.6. 0.6. This, contrarycontrary to tothe the Mach Mach flow flow plots, plots, would would suggest suggest that a stoichiometricthat a stoichiometric mixture wouldmixture be preferablewould be topreferable a lean fuel-air to a lean mixture. fuel-air The mixture. disparity The between dispar theity Machbetween and the Velocity Machplots and Velocity are due to plots the localare due speed to the local changing throughout the combustion chamber caused by the temperature increase, as seen in the next section.

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(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 Figureof 0.6, ( 14.c) Velocity (a) Velocity contour contourcontour for for equivalencefor equivalence equivalence ratio ratio ratio of of0.8, of 0.4, 0.4, (d ()b (Velocity)b Velocity) Velocity contour contour contour for for forequivalence equivalence equivalence ratio ratio ratio of 0.6,1.0.of 0.6, (c) ( Velocityc) Velocity contour contour for equivalencefor equivalence ratio ratio of 0.8, of (0.8,d) Velocity (d) Velocity contour contour for equivalence for equivalence ratio ofratio 1.0. of 1.0.

Figure 15. Variation of flowflow velocity with axial distance. Figure 15. Variation of flow velocity with axial distance. 4.2.3. Temperature 4.2.3. Temperature As expected, the combustion process releases energyenergy in the form of heat proportionally to the amountAs expected, of fuel introduced the combustion into the process combustion.combustion. releases Fi Figure gureenergy 16 inshows the form the staticof heat temperature proportionally of the to fuelthe throughoutamount of fuel thethe domain.domain.introduced TheThe into mixture mixture the withcombustion. with EQ EQ of of 0.4 0. Fi only4gure only reaches 16reaches shows about about the twice statictwice the temperaturethe temperature temperature of within the within fuel the combustor.thethroughout combustor. Thethe temperaturedomain.The temperature The for mixture the for stoichiometric thewith stoichiometr EQ of 0. mixture,4 onlyic mixture, reaches EQ 1.0, about increasesEQ 1.0, twice increases to approximatelythe temperature to approximately3.3 within times within3.3the timescombustor. the within constant Thethe constanttemperature diameter diameter section for the wheresection stoichiometr itwhere peaks itic atpeaks mixture, a ratio at a of ratioEQ 4.3 1.0, of in 4.3 increases the in throat. the throat. to Comparingapproximately Comparing the four-dithe3.3 timesfour-differentfferent within fuel-air the fuel-air constant mixtures mixtures diameter (Figure (Figure section 17a–d), 17a–d), where there there it is peaks only is only temperatureat atemperature ratio of 4.3 increase in increase the throat. within within Comparing the the free-jet. free- Thejet.the Thefour-different recirculation recirculation fuel-air zone zone remains mixtures remains at a(Figureat relativelya relatively 17a–d), low low there temperature temperature is only temperature of of 1500 1500 K Kcompared increasecompared withinto to thethe the 50005000 free- K observedjet. The recirculation by the NASA zone study. remains The at low a relatively temperatures low temperature at the nozzle of 1500part K may compared be attributed to the 5000 to the K presenceobserved ofby reverse the NASA flowflow study. at the pressureThe low outlettemperatures at the lower at the equivalence nozzle part ratios may of be 0.4 attributed and 0.6 at to least, the butpresence further of workreverse is clearlyflow at required the pressure inin thisthis outlet area.area. at the lower equivalence ratios of 0.4 and 0.6 at least, but furtherWith EQ work 0.4, is most clearly of therequired thermal in energythis area. release occurs in the divergent part of the combustion chamber.With TheEQ 0.4,energy most release of the ends thermal approximately energy release halfway occurs down in the the divergent combustion part chamber,of the combustion meaning thatchamber. all the The fuel energy is consumed release ends prior approximately to reaching the halfway throat. down With the the combustion high fuel-to-air chamber, mixtures, meaning the thermalthat all theenergy fuel isrelease consumed occurs prior from to injectionreaching untithe lthroat. the throat With ofthe the high combustion fuel-to-air chamber.mixtures, Thethe temperaturethermal energy in the release inlet canoccurs be observed from injection to be relatively until the low throat in all of four the cases. combustion chamber. The temperature in the inlet can be observed to be relatively low in all four cases.

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FigureFigure 16. VariationVariation of of temperature temperature with with axial axial distance. distance.

((aa)) (b)( b)

(c) (d) (c) (d) FigureFigure 17. 17.(a )(a Temperature) Temperature contour contour for equivalencefor equivalence ratio ofratio 0.4, of (b )0.4, temperature (b) temperature contour forcontour equivalence for Figure 17. (a) Temperature contour for equivalence ratio of 0.4, (b) temperature contour for ratioequivalence of 0.6, (c )ratio temperature of 0.6, (c) contourtemperature for equivalence contour for ratioequivalence of 0.8, ( dratio) temperature of 0.8, (d) temperature contour for equivalence contour equivalence ratio of 0.6, (c) temperature contour for equivalence ratio of 0.8, (d) temperature contour ratiofor equivalence of 1.0. ratio of 1.0. for equivalence ratio of 1.0. 4.2.4. Pressure 4.2.4.With Pressure EQ 0.4, most of the thermal energy release occurs in the divergent part of the combustion chamber.Figures The 18a–d energy and release 19 plot ends the approximately pressure drops halfway throughout down the the combustion combustion chamber. chamber, Thesemeaning results that Figures 18a–d and 19 plot the pressure drops throughout the combustion chamber. These results alldisplay the fuel the is loss consumed in performance prior to reachingof the engine. the throat. The pressure With the does high show fuel-to-air some fluctuation mixtures, thenear thermal the display the loss in performance of the engine. The pressure does show some fluctuation near the energythroat. release The slight occurs pressure from increase injection in until the throat the throat followed of the by combustiona pressure drop chamber. in the nozzle The temperature is expected in throat.as the Theflow slight is slowed pressure down increase before re-accelerating in the throat fo inllowed the nozzle. by a Itpressure is expected drop to in have the anozzle pressure is expected drop the inlet can be observed to be relatively low in all four cases. asacross the flow the iscombustion slowed down chamber before due re-accelerating to the formation in theof shock nozzle. trains It is within expected it. The to have presence a pressure of shock drop 4.2.4.acrossformation Pressure the combustion can even promote chamber mixing; due to however,the formation the trade-offof shock oftrains efficient within mixing it. The comes presence with of the shock formationdecrease ofcan total even pressure promote recovery. 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-off of efficient mixing comes with the decrease (a) (b) of total pressure recovery. (a) (b)

(c) (d)

(c) (d)

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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 formationAerospace 2019 can, 6, 135 even promote mixing; however, the trade-off of efficient mixing comes with16 ofthe 18 decrease of total pressure recovery.

(a) (b)

Aerospace 2020, 7, x FOR PEER REVIEW 16 of 18 (c) (d)

FigureFigure 18.18. (a) Pressure contourcontour forfor equivalenceequivalence ratio of 0.4, (b)) pressurepressure contourcontour forfor equivalenceequivalence ratioratio ofof 0.6,0.6, ((cc)) pressurepressure contourcontour forfor equivalenceequivalence ratioratio ofof 0.8,0.8, (d) pressure contour for equivalence ratio of 1.0.

Figure 19. Variation ofof pressurepressure withwith axialaxial distance.distance.

5.5. Conclusions TheThe aimaim ofof thisthis reportreport waswas toto adaptadapt thethe Dual-ModeDual-Mode Free-JetFree-Jet design, firstfirst tested by Trefny andand DippoldDippold III,III, andand fitfit thethe FAME18 FAME18 project project propulsion propulsion requirements requirements at at cruise cruise velocity velocity Mach Mach 8. The8. The goal goal of theof the FAME FAME project project was towas design to design an advanced an advanced engine configuration engine configuration for sustained for commercialsustained commercial hypersonic flight.hypersonic The study flight. was The based study on was the based prior workon the of prio ther FAME work teamof the to FAME provide team improved to provide designs improved of the inletdesigns and of nozzle the inlet for both and the nozzle ramjet for and both scramjet, the ramj whilstet and also investigatingscramjet, whilst the feasibilityalso investigating of the use the of afeasibility Dual-Mode of theFree-Jet use of combustion a Dual-Mode chamber Free-Jet design comb proposedustion chamber by Trefny design and Dippold.proposed The by dual-modeTrefny and combustionDippold. The engine dual-mode or dual combustion ramjet-scramjet engine is or a systemdual ramjet-scramjet that operates asis a system ramjet atthat low operates supersonic as a speedsramjet at and low then supersonic transforms speeds to a and scramjet then attransforms higher Mach to a scramjet numbers at between higher Mach 6 to 8. numbers This design between also benefits6 to 8. This from design being useablealso benefits over afrom wide being range useable of velocities over whilea wide keeping range of reasonable velocities performances. while keeping reasonableThe 2010 performances. numerical investigation by Trefny and Dippold III, was limited to hydrocarbon fuel operationThe 2010 and challengesnumerical wereinvestigation encountered by Trefny with extremeand Dippold static temperaturesIII, was limited within to hydrocarbon the combustion fuel chambers.operation and This challenges report adapted were the encountered Dual-Mode with Free-Jet extreme design static to fit temperatures the FAME propulsion within the requirements combustion atchambers. cruise velocity This Machreport 8 whileadapted using the hydrogen Dual-Mode instead Free-Jet of hydrocarbon-based design to fit the fuels. FAME propulsion requirementsIn this Computational at cruise velocity Fluid Mach Dynamics 8 while (CFD) using investigation hydrogen instead the k-ε ofmodel hydrocarbon-based was employed. fuels. The Y + valueIn was this maintained Computational at less Fluid than 1.Dynamics A Species (CFD) Transport investigation Laminar Finitethe k- Rateε model model was was employed. employed The to modelY+ value the was volumetric maintained reaction at combustionless than 1. process.A Species The Transport chemical mechanismsLaminar Finite were Rate integrated model intowas theemployed Fluent ANSYSto model software the volumetric package reaction using the comb CHEMKINustion process. tool (used The for chemical solving complexmechanisms chemical were kineticsintegrated problems). into the Fluent ANSYS software package using the CHEMKIN tool (used for solving complexThe Dual-Modechemical kinetics Free-Jet problems). design was validated against numerical results provided by NASA at MachThe 8. TheDual-Mode model captured Free-Jet welldesign all ofwas the validated salient flow agai featuresnst numerical such as results flow decelerationprovided by (although NASA at notMach quite 8. The as dramatically model captured as in well the NASAall of the model) salient and flow the features stagnating such recirculation as flow deceleration zone separated (although by 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 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, 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. 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 effort and subsequent discussion. Funding: This research received no external funding. Conflicts of Interest: The authors declare no conflict of interest.

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