Downloaded by guest on September 30, 2021 biu detonations oblique systems. power enabling advanced technology and propulsion detonation hypersonic ultra-high-speed develop to integrate pathway possible and a for allows breakthrough analysis experimental This diag- in simulations. to cor- Flow comparison are air. and through stabilization roborated and and is initiation hydrogen detonation chemiluminescence ramp, show and of imaging, a shadowgraph flow high-speed on as hypersonic such stabilized nostics, a wave, in detonation created oblique power. a reaction standing the space, present harnessing in A and fixed controlling we is for crucial that Here is detonation which setting. a high-enthalpy produces laboratory hypersonic that facility a a reaction abil- configuration, in our experimental them on unique study depends to and to igni- critical potential ity is propulsive the waves their detonation of compared of harnessing power understanding stability and propulsion An propagation, because greater tion, systems. interest for deflagration-based considerable use to of potential their now of are Detonation-based systems supernovas. and propulsion associated explosions generally blast reaction intense extremely of with ultrahigh- an mode are energetic are are highly which Detonations which efficient, waves. deflagrations, detonations, reaction be supersonic or speed may waves, reaction waves com- means therefore subsonic Such which and propulsion, reactions waves. exothermic rely nuclear bustion of will or forms likely chemical various we flight. of using propellants, reli- form supersonic as having some fuels and on on of hypersonic availability hinges the for robust, part, Given systems of large way propulsion in by able This, atmospheres planetary means. 2021) in controllable 5, February reentry review for and high-speed (received require 2021 flight will 31, travel March interplanetary approved and and terrestrial CA, Future Berkeley, California, of University Bell, T. Alexis by Edited and a Rosato A. propulsion Daniel hypersonic for detonation Stabilized PNAS combustion gain pressure hmclratos hs ae rvla aytmstespeed coupled the closely times many from at release travel energy waves by These driven reactions. chemical wave shock a of existing over advantage systems. technological power and this combustion propulsion Research to develop help (6). enhanced and will engines systems realize provide detonation detonation ultrahigh-speed to in rotating advancement ram detonation- shown where like benefits, have efficiency applications thermodynamic cycles for additional Even based no 5). are (4, there traditional cycles to compared based as 20%) deflagration to (∼10 scheme efficiencies thermody- cycle innovative increases namic considerably scram- an that is ramjets, propulsion Detonation hypersonic engines, for accelerators. jet ram and afterburning include jets, systems engines, These turbine (3). systems gas power and high- propulsion of superiority speed technological the maintaining for technology as routine as (2). travel today concepts, is intercontinental new travel mak- and intercity of in exploration set role space important One an ing (1). play vehicles could the engines, detonation-based drive modes to propulsion needed energetic highly are goal, ultimate this achieve To A combustion shock-induced n aemA Ahmed A. Kareem and rplinadEeg eerhLbrtr,Dprmn fMcaia n eopc niern,Uiest fCnrlFoia rad,F 32816; FL Orlando, Florida, Central of University Engineering, Aerospace and Mechanical of Department Laboratory, Research Energy and Propulsion eoaini uesnccmuto aeta consists that wave combustion supersonic a is detonation A transformational a are systems propulsion Detonation-based b aa etrfrSaeTcnlg,U aa eerhLbrtr,Wsigo,D 20375 DC Washington, Laboratory, Research Naval US Technology, Space for Center Naval heighg-pe ih tsproi n hypersonic focus. international an and and priority supersonic national a at now is speeds flight high-speed chieving 01Vl 1 o 0e2102244118 20 No. 118 Vol. 2021 | a yesncpropulsion hypersonic ao Thornton Mason , | hc-ae ecigflows reacting shock-laden a,1 | a oahnSosa Jonathan , | b hita Bachman Christian , doi:10.1073/pnas.2102244118/-/DCSupplemental at online information supporting contains article This 1 ulse a 0 2021. 10, May Published BY-NC-ND) (CC 4.0 NoDerivatives License distributed under is article access open This Submission.y Direct PNAS a is article This interest.y competing no the declare wrote authors K.A.A. The and tools; J.S., D.A.R., reagents/analytic and new data; analyzed contributed K.A.A. paper.y G.B.G. and G.B.G., and C.B., C.B. J.S., D.A.R., research; M.T., D.A.R., performed research; designed J.S. K.A.A. and and G.B.G., C.B., J.S., D.A.R., contributions: Author eoaini neprmna aiiyta rdcsrealistic engine. actual produces an in that use for adapted facility be can experimental which conditions flight an stabilized in a achieving in detonation difficulty problem historical the the by advances, compounded these is Despite (19–25). in numeri- stabilization concepts and their underlying important initiation, elucidated detonation have of simulations modes cal of number a shown high- the these in of expected waves concepts. be engine these would that of detonation conditions these high-enthalpy stabilization for speed, developing mechanisms robust in reliable and challenge finding computa- initiation The is and study. concepts experimental this engine the ODWE of to an results relation by tional the powered illustrates vehicle 1 Fig. hypersonic and vehicles. conceptual launch a space needed reusable shows speeds theo- other the its and to for spaceplanes aircraft here for hypersonic interest propel particular to of ability is retical (3, ODWE (ODWE) The engines 16–18). wave 7, detonation oblique (13–15), and engines standing detonation and rotating det- 10–12), engine pulse (5, attention: detonation engines research onation of significant categories received concept main have this concepts Three of implementation difficult. the been although has (3), it new genera- time not energy average is and propulsion tion the The for waves journey. than detonation same using New faster the of complete idea from times to 6 flight Concorde 5 of legendary half-hour is the number a took and to Mach London of comparable flight to is case vehicle York That the a Mach (7–9). in a to 17 with as corresponds to operating path 5, engine flow Mach An 5 of mixture. fuel speeds hydrogen–air reaching a often sound, of owo orsodnemyb drse.Eal [email protected] Email: addressed. be may correspondence whom To hthsteptnilt eouinz ihsedpropulsion high-speed future. revolutionize the of to potential phenomenon the a has detonation, that oblique conditions stabilized dis- flow a the and generate goal: configuration that this experimental attaining an in of This step covery significant emissions. a com- reduced reports thermodynamic of work and higher reliability, form provide enhanced powerful to efficiency, potential most the such the has basing detonations, bustion, of on exit possibility system and The entry a atmo- efficient atmospheres. our allow planetary through and supersonic speeds from flight high allow and very would at hypersonic sphere system a for develop Such to systems flight. effort international propulsion intensifying robust an now is There Significance aoaoyeprmnsadnmrclsmltoshave simulations numerical and experiments Laboratory b are .Goodwin B. 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ENGINEERING ing at Mach numbers ranging from 6 to 17, largely dependent upon engine inlet design (7–9). The fueled case shown here has a mixture molar composition of major species H2/O2/N2/H2O = 13.2/9.3/62.0/14.7% (yielding a global H2/O2 equivalence ratio of φTS = 0.71). Prior to fueling the facility, the nonreacting flow field was ana- lyzed to confirm the oblique shock wave produced by the ramp matched the theoretical adiabatic oblique shock solution for a 30◦ ramp. For the given nozzle area ratio (A/A∗ = 25), the non- reacting hypersonic flow shows the predicted oblique shock angle (β) of 42◦ for an inflow Mach number of 4.4 with a ratio of specific heats (γ) of 1.3. Once fuel is introduced, an ODW is Fig. 1. Schematic of oblique detonation engine concept. The experimental initiated over the ramp and is sustained for the duration of the and computational ODW domains are highlighted along with their location experimental test, approximately 3 s. During the reaction, the in the engine flow path. highest chemiluminescence signal intensity is observed imme- diately above the ramp due to the presence of the detonation wave at that location. The sustained detonation is shown by the Previous experimental studies were unable to show a stabilized reacting shock structure (RS2) in Fig. 2C. As the incoming flow oblique detonation wave (ODW) for an extended period, due to passes through S2, it enters the induction region. In the induc- their use of shock/expansion tubes or projectiles (7, 22, 26–28). tion region, the mixture is heated by the temperature rise across These types of facilities have limited run times, on the order of the shock. This heating allows for the reaction process to occur microseconds or milliseconds. Another major difficulty in stabi- through autoignition and the formation of a detonation wave lizing the detonation wave is upstream wave propagation through with a steeper angle, RS2 (73◦) (31). The flow velocity is cal- the boundary layer leading to unstart with recent experiments culated as being 99.7% of the theoretical detonation wave speed showing deflagration-to-detonation transition in a hypersonic for a freely propagating normal detonation in this mixture, UCJ . flow and an unstable detonation that propagated upstream (24). The static pressure profile shown in Fig. 3D, measured down- Several numerical studies have shown potentially steady ODW stream of the ramp, shows a clear pressure rise generated by the but lack experimental verification (21, 23, 29, 30). These leave reaction when compared to the baseline nonreacting pressure uncertainty about the stability of ODW, which must be addressed trace over the duration of the test without activating the fuel. The through experiments capable of creating the appropriate peak pressure reaches 2.7 times the baseline nonreacting pres- conditions and maintaining them for an extended period. sure and 10.5 times the nozzle exit pressure. The velocity balance This paper reports results from a study demonstrating exper- and the pressure rise measurements are strong conformation of imentally controlled detonation initiation and stabilization in a the detonation formation. hypersonic flow for a situation similar to proposed flight condi- tions for these vehicle concepts with an active run time of several Mechanism of the Oblique Detonation seconds. The experimental results capture the stabilized deto- An ODW is sustained for the duration of active fueling. Fig. 3 nation, as shown in the shadowgraph and chemiluminescence shows a sequence of images along with the pressure trace for images, and are further confirmed and explained by the theory and numerical simulations of the system. A 30◦ angle ramp is used in the high-enthalpy hypersonic reaction facility to ignite and stabilize an ODW, shown schematically in Fig. 2A. The shock-laden, high-Mach number flow induces a temperature rise to ignite and stabilize a detonation in the incoming hydrogen–air mixture. The combination of matching the flow Mach number to the MCJ conditions and low boundary layer fueling result in the stabilized detonation. Static pressure measurements confirm a pressure rise induced by the detonation wave. High-fidelity computational fluid dynamics simulations have been used to pro- vide additional detailed insight into the detonation initiation and stabilization process. Stabilizing Detonations in a Hypersonic Flow A detonation is stabilized on a ramp in a hypersonic flow as shown in Fig. 2. The images in the figure show the flow den- sity gradients (shadowgraph) with the chemiluminescence from the chemical reactions overlaid. Fig. 2B shows the baseline non- reacting hypersonic flow in which the preburner was operating and the main fuel injection was not activated leading to no addi- tional chemical reactions in the test section. Fig. 2C shows the same hypersonic flow with the fuel turned on, which resulted in the generation of a stabilized ODW. The hypersonic flow is produced by an axisymmetric Mach 5 converging–diverging noz- zle as shown in Fig. 2A. The fuel and air are premixed slightly upstream of the nozzle throat, detailed in Materials and Methods. The turning angle of the ramp is θ = 30◦. The flow stagnation pressure (P0) is 5.63 MPa, and the stagnation temperature (T0) is 1,060 K, resulting in an effective exit Mach number of 4.4, a value expected within the engine flow path of a vehicle fly- Fig. 2. (A) HyperReact. (B) Nonreacting flow field and (C) stabilized ODW.

2 of 7 | PNAS Rosato et al. https://doi.org/10.1073/pnas.2102244118 Stabilized detonation for hypersonic propulsion Downloaded by guest on September 30, 2021 Downloaded by guest on September 30, 2021 ecigts eto ttcpesr ai pesr fratn case reacting of (pressure ratio pressure static section [P test reacting (D) 3. Fig. tblzddtnto o yesncpropulsion hypersonic for detonation Stabilized cycle-to-cycle al. et the Rosato a to through due goes detonation reaction downstream. along underdriven-to-overdriven the and reactions of that upstream the variation believed travel inflection while is cyclically the RS2 It at surface and remains ramp S2 front the shocks reaction series leading between time image The point shadowgraph 3. the Fig. the in in in seen fluctuations as the the front to detonation shock While responds The fashion. dynamically reaction. cyclical ahead a the structure in fluc- run of front the detonation throughout duration the slightly of tuates the sur- location the the for sustained, above is ramp detonation remained the front of detonation face The case. reacting a R /rsueo oratn ae[P case nonreacting of ]/pressure (A–C h eoainsrcuefrtresae uigrnand run during stages three for structure detonation The ) NR )v.time. vs. ]) n oa ndrcin farw hw.Sldlnsrpeetcniin for conditions represent lines Solid shown. arrows φ of directions in polar ing test the times 0.30 to wall the from and height section test ledaodmre ersnstets conditions. test the represents marker diamond Blue 4. Fig. (φ ratios in discussed section, priate test the in in detail profile more fuel the of con- measurement be H to through accomplished needs was premixing This of sidered. level the composition, mixture the and the no composition Because have freely value solution. same that stable would below the numbers detonation of Mach Flow a mixture temperature. static which quiescent polar at a in shock number propagate Mach The the ramp. number, is Mach the (CJ) which Jouguet to Chapman attached the from remain originates will ODW the and calculated stabilized, detonation the be which can for angle in turning shown minimum polar, the is shock the on between is 4, exist band Fig. that stability conditions The the detonation. as the possible defined by is amount produced and release stability temperature, heat ODW static of angles composition, which mixture turning given over of a numbers for range model Mach the The flow of (20). and estimate the achieved be theoretical predict can a to stability generates used ODW is which model at and limits flow low downstream and receding compressible a reaction A overdriven Conversely, the deflagrating. in is flow. result that will the release detonation to high heat a A opposing in flow. heat upstream, number result propagates and high-Mach will composition the release in mixture heat reaction in the for balance release ideal the images. these achieving in seen not is broad section the test in from the occur luminescence in species species Hence, broad range. the wavelength while UV range the that wavelength emissions visible luminescence in strongest leading are the highlight the signal to chemiluminescence above The filtered wall. is front, top the at reaction takes and front, denotative burning reaction Additional the flow. behind reacting place the of nature turbulent

TSL θ

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ENGINEERING section height were calculated. These heights correspond to the A full height of the ramp and the approximate height at which the reaction and triple point form and were chosen to encom- pass the fuel most likely to pass through the induction region and detonation wave. For the lower segment, φTSL AVG was cal- culated as being 0.24, and for the upper segment, φTSL AVG = 0.44, resulting in local MCJ values of 2.95 and 3.68, respec- tively. Fig. 4 shows the ODW stability limits with the range for both values highlighted in the red shaded area. The flow Mach number and the ramp angle θ for the experiment do not vary. This places the test condition (M = 4.4, θ = 30◦) within the the- oretical stability limits created by the shock polars for these conditions. Numerical Simulations B Numerical simulations were performed at conditions approxi- mating those achieved within the experimental facility in order to corroborate the experimental results and compare, in particu- lar, the structure of the ODW. The simulations solve the reacting Navier–Stokes equations for a compressible fluid by numerical integration with fifth-order accuracy in space and second-order in time on a Cartesian, dynamically adapting computational grid. This method is discussed in detail in ref. 32. The maxi- mum computational cell size is 1.4 mm, and the minimum is 11 µm. The reactions are modeled with a simplified, calibrated chemical-diffusive model (CDM) that uses a single Arrhenius reaction rate to convert reactants to products. The CDM has been used extensively in detonation studies, shown to reproduce desired combustion properties, such as detonation wave speed and flame temperature (33–35), and was used recently to study ODW ignition and stability characteristics when a boundary layer is present on the wedge surface (32). For this study, the CDM was optimized for a hydrogen–air mixture. The supersonic reactive flow over the ramp was modeled in an idealized way, in a rectangular domain with a diagonal inflow from the left and right boundaries, using a boundary condition on the lower wall to model the flow interaction with the ramp. Thus, the domain and conditions were constructed as if rotated 30◦ in order to model the ramp angle on an orthogonal mesh, shown in Fig. 5A. The visualization of the results in Fig. 5B has been rotated to the experimental frame of reference and cropped to more accurately represent the experiments. C A snapshot of the numerical result is shown in Fig. 5B, and the same image is overlaid on the experimental shadowgraph in Fig. 5C with the relevant structures aligned. The lack of turbu- lence present in the simulations compared to its abundance in the experiments necessitated an attempt to make up for local compressibility effects and temperature fluctuations in the exper- iments by simulating a higher static temperature inflow. This is seen by the minimum temperature bound on the color map in Fig. 5B, which is higher than the static temperature of the fuel–air mixture in the experiments, and was required for ODW initiation in the simulations. In this way, the enthalpy of the incoming fuel–air mixture in the simulations is enhanced beyond that of the experimental facility in order to compensate for the inability of the simulations to reproduce the experiment’s turbu- lence, which has been shown to assist in mixture ignition through Fig. 5. (A) Computational domain overlaid on test section ramp geometry. the generation of eddy shocklets and local compressibility effects (B) Simulation results showing the ODW temperature field. (C) Experimental (36). The inflow Mach number is 5 (Mach number generated in shadowgraph image of the ODW structure overlaid with simulation results the experimental facility after full expansion of an incoming air- from B. flow), and the static pressure matches that of the experiments. A burning boundary layer is present due to the great amount of viscous heating that occurs along the no-slip boundary that which propagates an ODW. It is seen in Fig. 5C that the lead- is imposed on the ramp surface. Above the boundary layer, the ing oblique shock wave, triple point, and ODW are evident in freestream passes through an induction region in which the flow both experiment and simulation, all of which are important char- autoignites, forming a reaction front which steepens and inter- acteristics of the traditional ODW structure. In the simulations, sects the leading oblique shock wave. This intersection forms a there is clear detonation cell structure, typical of a propagating triple point of extremely high pressure and temperature, out of detonation wave.

4 of 7 | PNAS Rosato et al. https://doi.org/10.1073/pnas.2102244118 Stabilized detonation for hypersonic propulsion Downloaded by guest on September 30, 2021 Downloaded by guest on September 30, 2021 P)adttltmeaue ,5 o110K tbeoblique stable A K. 1,100 to 1,050 temperatures total and or MPa) extinguish either to downstream cycle. back the recedes repeat CJ noz- then the the and of enters reaction zle 80% shock-coupled exceeds The reactions velocity. detonation these The recov- of release upstream. temperature velocity propagating heat detonation propagation rapid high overdriven more with an disk. and to shock Mach rates leading normal reaction a higher a forms enabling is and ery, ramp disk the Mach by A generated the shock the with for- intersects oblique propagate on wave and reaction point is pressure forward-propagating build reaction initiation The to ward. begins the the reaction II, from the wall, regime Starting far in nature. cases in the deflagrated oscillatory For 6, occurs. Fig. combustion in surface. I ramp the regime over experienced by equiv- are reactions and represented pressures, ratios, total temperatures, varied. alence total were low composition relatively mixture At and turning temperature, flow stagnation The at Fig. facility. constant conditions. held the was using of angle tested range conditions of broad the behaviors controllability shows a 6 the over reaction modes and of burning evolution of forms different the process showing major paper. the observed three this were in in process, explored highlighted this waves were During detonation conditions stable the of pursuing range large A Regimes Flow Reacting Hypersonic regime and combustion; shock-induced disk Mach II, regime combustion; shock-induced oblique I, regime ODW. operation: III, of modes major of chemiluminescence i.6. Fig. tblzddtnto o yesncpropulsion hypersonic for detonation Stabilized al. et Rosato eieIIocr ttehgettse rsue 56t 5.9 to (5.6 pressures tested highest the at occurs III Regime shock-induced increased, are pressure and temperature the As (Left prtn odtostse ihrm for ramp with tested conditions Operating ) θ 30 = ◦ hl h tgainpressure, stagnation the while θ = 30 ◦ ihsaiiympfrteratn etcniin.(Right conditions. test reacting the for map stability with h eprtrsadflwMc ubr fti facility. at this ODW of cases, stable numbers a other Mach establishing flow all in and to factor temperatures critical the compared the illustrate as be to stag- regime, to additional appears this used The in case III. pressure regime The within nation 1.2. falls ODW to sustained 0.7 the approximately at from section ing test the within observed is detonation I vrg rsuertov.time. vs. ratio pressure average III 7. Fig. A Peak Pressure Ratio (PR/PNR) 2.60 2.70 2.80 2.65 2.75 0 0 0010 1200 1100 1000 900 800 . . . . 1.2 1.1 1.0 0.9 0.8 Regime (B) III). (regime cases ODW for rise pressure Normalized (A) Stagnation Temperature[K] φ TS https://doi.org/10.1073/pnas.2102244118 Pressure Ratio (PR/PNR) 1.0 1.5 3.0 0.0 2.0 0.5 2.5 B 0 2.4 2.6 vradshadowgraph- Overlaid ) 59 24 . 6 Time [s] Stable 5 ODW φ . TS 7 PNAS ausrang- values 6 5 . 8 80 5 | . f7 of 5 1

ENGINEERING Static Pressure Transducer Main Fuel Co-Flow Air (x4) Pre-Burner Ts = 280 K T0 = 1060 K Pre-Burner Air θ = 30° P0 = 5.70 MPa M = 4.4

45 mm Pre-Burner Fuel

Mixing Chamber CD Nozzle Test Section 350 mm 219 mm 159 mm

Fig. 8. Schematic of HyperReact experimental facility.

Reactions in this regime cause a pressure rise within the test of O2 and H2 in the flow at those locations, and the mole fraction of the section, as compared to the nonreacting baseline case at similar additional species found is provided in the format (%H2/%O2/%N2/%H2O). total pressure and temperature. Fig. 7A shows the ratio of the The premixing level of the fuel results in a test section fuel profile that peak test section static pressure during regime III reaction cases is shown in Fig. 9, which was experimentally determined through Raman with respect to the baseline static pressure. Across regime III, spectroscopy measurements during nonreacting operation of the local H2 concentration. The local premixed mixture equivalence ratio (φ ) near the the pressure rise remains at approximately 2.7 times greater than TSL ramp surface is then used in calculating φTSL AVG, defined as being the aver- the baseline case. Fig. 7B shows the average pressure ratio of all age fuel concentration between the test section wall at y/h = 0 and the measured regime III cases over the run time. The profile shows selected upper boundary. ODW characteristics, including MCJ and ODW sta- a consistent and repeatable pressure rise once the hydrogen fuel bility limits, were calculated using the φTSL AVG values determined by this is injected. method. A 30◦ turning angle ramp is used for stabilizing the detonation wave. Materials and Methods The ramp spanned the full width of test section and is placed at 44 mm downstream of the CD exit plane. The height of the ramp is fixed at 7.5 mm The High-Enthalpy Hypersonic Reacting Facility (HyperReact) at the Univer- to avoid a blockage ratio higher than 17% within the test section. The aft sity of Central Florida (UCF) is used for this study as shown in Fig. 8. The face of the ramp is relieved at a 3◦ angle relative to the test section wall. facility consists of five major components which are, in the order of their This allows the flow to partially reexpand along its length. Test section static location along the axial direction of the facility, an in-flow preheater, mixing pressure measurements are acquired at the test section’s top wall midplane, chamber, main fuel injection stage, converging–diverging (CD) nozzle, and marked with a red dot in Fig. 8. optically accessible test section. The in-flow preheater consists of a coaxial The ODW is recorded using simultaneous high-speed schlieren and vis- hydrogen–air jet flame surrounded by evenly spaced coflow air jets consum- ible range wavelength 450 to 875 nm chemiluminescence imaging. The ing 44% of the oxygen. The preheater is controlled to achieve a stagnation test section has fused quartz windows on the side walls for full optical temperature range of 800 to 1,200 K, corresponding to a static temperature access to an interrogation region of 105 mm long and 45 mm high. The of 180 to 320 K in the test section. The mixing chamber consists of a square schlieren system consists of a Z-type setup using two 152.4-mm spherical channel with an internal height of 45 mm and a length of 350 mm. This mirrors, with focal lengths of 1.52 m, and a high-power Luminus PT-121-G segment of the facility allows for homogeneous mixture in-flow feed to the LED light source. Both the schlieren and chemiluminescence images are cap- CD nozzle. The main fuel injection used for the downstream reactions intro- tured using Photron SA1.1 high-speed cameras recording at 30 kiloframes duces the supplementary fuel prior to entering the CD nozzle to allow for premixing. The CD nozzle has an axisymmetric square cross-section along the entire length of the nozzle. The characteristic length scale for the noz- zle is the 45-mm height for both the inlet and exit, and the throat height is 1.0 9 mm. The inlet-to-throat and exit-to-throat area ratios are both 25:1. The contracting section of the CD nozzle is designed to produce a uniform veloc- ity profile at the throat and minimize boundary layer growth as detailed by Bell and Mehta (37). The diverging section of the nozzle consists of a three- 0.8 dimensional contour derived from an analytical method by Foelsch (38), and a cubic matching function is used (39) to smoothly transition between the two segments of the nozzle. Additional details on the nozzle design can be found in ref. 40. The CD nozzle is designed to provide an exit Mach num- 0.6 Laser Measurement ber of M = 5.0 for dry air at 300 K (24, 40). The effective Mach number is Location dependent on the temperature and composition-dependent heat capacity

ratio of the mixture entering the nozzle for the test, which results in a range y/h of 4.3 to 4.6. The CD nozzle issues the hypersonic flow mixture to the opti- 0.4 cally accessible test section consisting of a square channel of height 45 mm and length 159 mm. The fuel used for the preheater stage and the main fuel injection is 99.99% ultrahigh-purity hydrogen. Air is provided from a pressure source tank at 34.45 MPa. 0.2 Fuel and air mass flow rates supplied to the facility are metered through precision choked orifices. The air orifice is 4.57 mm in diameter. The orifices for the preheater fuel and main fuel injection lines vary in size to accommo- date the broad range of fueling flow rates needed to cover the extent of 0.0 conditions tested. Fuel orifice sizes used range from 0.56 to 1.57 mm in diam- 0.0 0.5 1.0 1.5 2.0 eter depending on the mixture fraction. Pressures upstream of each choking orifice are measured using Dwyer 626 absolute pressure transducers with φTSL ranges of 0 to 20.68 MPa and accuracy of 1% of the full-scale range. The

equivalence ratios of both the preburner (φburner ) and the downstream con- Fig. 9. Schematic of fuel measurement location and curve-fitted local fuel ditions in the test section (φTS) are calculated based solely upon the amount concentration. Limits used to determine φTSL AVG are also shown.

6 of 7 | PNAS Rosato et al. https://doi.org/10.1073/pnas.2102244118 Stabilized detonation for hypersonic propulsion Downloaded by guest on September 30, 2021 Downloaded by guest on September 30, 2021 1 .Edla,“us eoainegn- ttsrve n ehooydevelopment technology and review status engine-a detonation “Pulse Eidelman, S. Chal- 11. propulsion: detonation Pulse Netzer, W. D. Borisov, A. A. Frolov, M. S. Roy, D. G. 10. 5 .Hsia .Fjwr,P oasi udmnaso oaigdetonations. rotating of Fundamentals Wolanski, P. Fujiwara, T. Hishida, challenges, M. Experimental propulsion: wave 15. detonation Rotating Braun, M. E. Lu, K. F. 14. report” status brief A concept: engine engines. detonation-wave detonation rotating “The pulse Kailasanath, of K. analysis 13. cycle Thermodynamic Pratt, T. D. Heiser, H. W. 12. 6 .Dbr,J-.Boa Sadn omldtntosadolqedtntosfor detonations oblique and detonations normal “Standing Broda, J.-C. Dabora, E. 16. 1 .Lu .S i,D u -.Wn,Srcueo nolqedtnto aeinduced wave detonation oblique an of Structure Wang, J-P. Wu, D. detonation Liu, Y.-S. oblique Liu, standing Y. of Morphology 21. Glenn, E. D. Humphrey, W. J. Pratt, T. trans- waves. D. detonation for Oblique combustors 20. Gross, Wave A. Bowles, V. R. J. Cambier, 19. J.-L. Adelman, G. H. Menees, P. G. 18. Wola P. 17. 2 .I ors .R ae,R .Hno,“hc-nue obsini high-speed in combustion “Shock-induced Hanson, K. R. Kamel, R. M. Morris, I. C. 22. aaAvailability Data a with 300 operated of was lens, f/1.2 mm 350 of 50 resolution Nikor Nikon mm a 300 with 640 equipped to 164 camera, a approximately 70 with of Nikon resolution images a spatial and a with lens equipped f/5.8 is to camera f/4 schlieren The second. per tblzddtnto o yesncpropulsion hypersonic for detonation Stabilized al. et Rosato .W .Sret .A rs,Dtnto aehproi ramjet. hypersonic wave Detonation Gross, A. R. Sargent, H. W. 9. investiga- experimental and “Analytical Cambier, J.-L. 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ARS IASieh2020 SciTech AIAA IAppendix SI 543–549 30, 545–672 30, IAJ. AIAA Shock 38, . 0 .Ss,K .Amd Dsg eeomn fahproi obso o oblique for combustor hypersonic a of development & “Design Ahmed, A. K. Sosa, J. 40. Heybey, H. W. Crown, C. J. 39. shocks. oblique behind structures Detonation Oran, S. E. Kailasanath, K. detonation Li, oblique C. “Wedge-stabilized 31. Ahmed, K. Goodwin, B. G. wedge: a Bachman, by wave L. detonation C. oblique an 30. of Stabilization Deshaies, B. Silva, Da F. F. L. 29. detonation oblique of “Onset Deshaies, B. Desbordes, D. Silva, Da F. F. L. Viguier, C. 28. Kasahara hypervelocity J. a 27. on stabilized detonation uni- “Oblique Shepherd, A E. Taylor, J. Kaneshige, D. J. B. M. Gamezo, 26. N. V. Ahmed, K. Chambers, wedge. J. confined Poludnenko, a Y. A. by 25. induced waves Detonation Wilson, Sosa J. R. D. 24. Fan, H. Lu, K. F. 23. 8 .Fesh h nltcldsg fa xal ymti aa ozefraprle and parallel a for nozzle laval symmetric axially an of design analytical The Foelsch, K. 38. tunnels” wind low-speed small for design “Contraction Mehta, D. R. Bell, tur- H. Compressible J. Gamezo, 37. N. V. Poludnenko, A. Ahmed, A. a K. in Chambers, detonation J. to Sosa, transition J. and stability 36. flame Premixed Oran, S. E. Goodwin, gas- in B. transition G. deflagration-to-detonation the 35. of Origins Gamezo, N. V. Oran, S. E. 34. lay- boundary of A. effect Kaplan, the R. C. and 33. criteria Ignition Goodwin, B. G. Bachman, L. C. 32. o r hita aha h efre h opttoa simulations Program. computational Base NRL the the performed by who provided was Bachman Support Fellowship. Christian Karles Mr. (NRL) was Laboratory for Sosa Research and Jonathan Naval Dr. the Fellowship for by Support provided Improvement Fellowship. Award Doctoral Dissertation Presidential NSF UCF the by the Space through Florida supported NASA Consortium are the by analyses Grant provided was The support student Li). Graduate 1914453. Chiping Pro- Dr. FA9550-19-1-0322, Manager and (FA9550-16-1-0441 gram Research Scientific of Office ACKNOWLEDGMENTS. eoainwv tblzto”in stabilization” wave detonation 1950). MD, Oak, White in freestream” premixed Forum 2019 hydrogen-air Energy hypersonic a in waves study. numerical parametric A hydrogen-air for results 3023–3031. numerical and in experimental mixtures” between Comparison waves: (2002). mixtures. acetylene/oxygen/krypton in bodies cal supernovae. 3015–3022. IA ter- in projectile” in type transition and deflagration-to-detonation systems (2019). unconfined chemical for restrial mechanism fied (2020). 7596–7606 38, Technol. Sci. Aerospace 07,p 0371. p. 2017), jet. uniform https://ntrs.nasa.gov/ 1988) Server, 2021. 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