Hypersonics: laying the road ahead

Pedro Goncalvespedro.simoes.goncalves@tecnico.ulisboa.pt

Instituto Superior Tecnico, Universidade de Lisboa, Portugal

July 2018

Abstract

This work provides a technology roadmap for hypersonic applications and use it to some extent inpreliminary hypersonic aircraft design. A survey on barriers preventing hypersonic flight was conductedand a technology roadmap was created based on novel concepts which could address the critical require-ments. Three engines (ramjet, scramjet and rocket) were numerically modeled in SUAVE, a frameworktool developed by Embraer and Stanford University, to run low-fidelity hypersonic vehicle analysis.Since the tool lacked a detailed hypersonic background, simple aerothermodynamic and weight distri-bution capabilities were also added. These new models were compared against numerical codes and/orexperimental data for validation before being implemented in SSTO and atmospheric cruise missionscenarios that assessed the usefulness of new technologies and the performance of its subcomponents.With this work, a new proposal for the future of hypersonics is presented and SUAVE is now able torun basic hypersonic simulations;Keywords: Hypersonic, Roadmap, Technology, Low-fidelity

1. Introduction

Imagine traveling from London to Sydney in lessthan four hours. Imagine a regular round-trip flightaboard a commercial spaceplane to a space hotelin Low Earth Orbit (LEO) for an affordable price.More than half a century after the technologicalmarvels that put Mankind on the Moon, these twoconcepts still seem somewhat eccentric and share acommon notion of uncertainty. As of March 2018,a flight from London to Sydney doesn’t take lessthan 17 hours (all-time record); on the other hand,the only handful of space tourists who have vis-ited and stayed at the International Space Stationsince the turn of the century have paid no less than$20,000,000 (2018) for their journey, far from af-fordable. However, all of these projects may be-come a reality if sustainable hypersonic capabilitiesare provided.

Hypersonic flight is the next major leap forwardin aerospace and despite not being an entirely newconcept, it has been gathering more attention fromthe scientific community in recent years. The abil-ity to significantly cut travel times and bring cities,nations and people even closer will revolutionizeand strengthen interpersonal and economical ties,as well as redefine the notion of distance: travel-ing to the opposite side of the world by plane couldtake almost the same time as a roadtrip across theentire extent of continental Portugal. Rightfully so,this justifies the research and development of newtechnologies to enable sustainable and affordable

hypersonic transport. The implications of a suc-cessful hypersonic program are significant, not onlyfor terrestrial applications but also for routine spaceaccess, a goal that dates back almost 60 years andis yet to be accomplished. Not only hypersonics canconnect people all over the planet, it can become abridge towards Space exploration and colonizationby lowering costs associated with Space endeavors.

Mastering hypersonics is a complex challenge andwhile this work cannot, in any way, achieve such anencompassing level, it serves as a starting point totackle the many intricacies of hypersonic flight; thisincludes understanding current technological barri-ers and gaps, map new enabling technologies withthe potential of unlocking hypersonic capabilitiesand apply them in a preliminary design analysistool, SUAVE, to better understand how they mayimpact overall aircraft performance.

2. Background2.1. Roadmap requirements

Before beginning to map enabling technologies inthe field of hypersonics, a roadmapping structureand strategy are required.

A technology roadmap is a tool to support strate-gic and long-range planning by matching both shortand long term goals to specific technological solu-tions. For the purpose of this work, the roadmapstructure used is adapted from the Defense LogisticsAgency of the United States Department of Defense[2].

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To evaluate each technology, a scale is requiredto provide a universal comparison method. Amongother scales, the technology readiness level (TRL)was selected as a reference scale for this project.By using a standard TRL, the common assessmentof individual technologies allows for risk reductionboth in budget and planning processes. For the pur-pose of this work, the same TRL standard as thatof the European Space Agency for Space productdevelopment is used [1], presented in Figure 1.

Figure 1: TRL scale [1]

2.2. State of the artUnderstanding hypersonic state-of-the-art isparamount to the success of this work. By do-ing so, it is possible not only to list hypersonicconcepts and technologies that have already beendemonstrated, but also to comprehend the techno-logical barriers preventing an extensive hypersonicapplication.

Sonic boom Currently there are no industrystandards for the acceptability of a sonic boomand the overground ban of supersonic commercialtravel from the 1970s still endures. NASA is cur-rently experiencing with design innovations such asboom shaping for the Quiet Supersonic Technologydemonstrator (QueSST), with the goal of reassess-ing current regulations and providing a quieter al-ternative for supersonic transport [15].

Aerodynamics One of the most promising con-cepts in hypersonic aerodynamics is the waveriderconfiguration which makes use of shock waves be-ing generated by the airframe to act as lifting sur-face, in a phenomenon known as compression lift.

The Boeing X-51 is the latest vehicle known to havedemonstrated such capability, in 2013. However,despite its success, there are major technological is-sues preventing a wider application. Scaling thisconcept to the size of commercial airliners or space-craft is currently unfeasible as a result of the verylow volumetric efficiency of the waverider concept.Another problem is that this design only works fora particular Mach number, resulting in a significantreduced lift capability when operating in off-designconditions.

Propulsion There are severe limitations in cur-rent propulsion systems that render sustained air-breathing hypersonic flight unachievable. Air-breathing engines outperform conventional rocketsin regards to specific impulse, safety and maneuver-ability. However, they are substantially harder todesign and test.

Conventional turbojet engines cannot single-handedly propel an aircraft to Mach 3+. On theother hand, supersonic and hypersonic propulsionconcepts such as ramjets and scramjets are onlynow becoming feasible but they require supersonicspeeds to be ignited. As a result, they are uselessduring the subsonic portion of flight. If a spaceplaneis envisioned, rocket engines must be incorporatedinto the propulsion system for orbital assist. How-ever, because oxidizer has to be carried on-board,there is a significant payload weight penalty. Thispenalty is, in fact, so substantial that current rocketengines cannot carry enough propellant efficientlyin a single stage to reach orbit; this drives the costsup as multiple stages are required. The most recenttake on this issue is SpaceX’s concept of reusablefirst-stage engines.

The most powerful, proven, reusable high-speedair-breathing engine conceived to date is the Pratt& Whitney J58, which powered the Lockheed SR-71 ”Blackbird” up to Mach 3 with technology devel-oped 60 years ago. This could only be possible witha high bypass afterburner at a very high fuel burnrate, effectively making it a turbo-ramjet engine.

Materials Regarding airframe and other aircraftexternal structures, sustained hypersonic flight isstrongly limited by the materials used. To sur-vive the harsh hypersonic conditions, they must bestrong enough to withstand high heat fluxes, tem-peratures and temperature gradients. They mustalso provide shielding against oxidation and erosionwhile supporting strong aerodynamic loads. Crit-ical conditions take place on the vehicle nose andwing leading edge. Temperature requirements on asharp leading edge for a cruise hypersonic vehicleare in the order of 2000 to 3000◦C, substantially

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higher when compared with the reinforced carbon-carbon used in the Space Shuttle orbiter.

TPS Thermal protection systems (TPS) incorpo-rate different technologies that help the vehicle re-sist excess heat in hypersonic flight. They are notonly dependent on the material properties but othermechanisms that can be implemented to alleviatethermal loads. State-of-the-art TPS has been ap-plied to re-entry spacecraft. When it comes toreusable concepts, the Space Shuttle TPS, com-posed of TUFI ceramic tiles, could sustain multiplere-entries at up to 100 W/cm2[17]. However, uponlanding, panels would require significant mainte-nance, resulting in high turnaround times and re-pair costs. Ultimately, mechanical failure (i.e., verylow impact resistance) was the main reason behindthe fatal Challenger disaster, showing how suscep-tible the system was to mechanical anomalies.

GNC One of the biggest accomplishments inGuidance, Navigation and Control (GNC) systemsdates back to the 80s, when then the Soviet Unionmanaged to launch and land Buran, a fully auto-matic spaceplane. This was an extraordinary featconsidering how rudimentary GNC theory and tech-nology was back in this age. Buran was the firstspace shuttle to perform an unmanned flight andstill holds the record of the largest aircraft to or-bit and land unmanned [3]. Recently, there hasbeen progress in the hypersonic GNC field, suchas demonstrations for the Boeing X-51 scramjet-powered aircraft. Nevertheless, one of the most ad-vanced GNC systems of today must be that of theBoeing X-37. This secretive spaceplane is taken toLEO through expendable rocket engines, where itthen conducts its classified missions before perform-ing a fully autonomous re-entry and landing

2.3. Novel conceptsTo fill in the gaps from the current state-of-the-art,a survey of enabling technologies was conducted.The most promising concepts were listed and de-scribed with a moderate level of detail to provide abasic overview of their functionality. Table 1 sortsthem by technological area and associates them toa TRL.

2.4. Roadmap implementationUpon reviewing some of the most significant tech-nological prospects for hypersonic applications, thenext step in the roadmapping procedure is to se-lect the best alternatives to fill in the requirements.This recommendation is based on the current TRL,the current R&D efforts towards its completion andexpected timeline for delivery. To do so, data onprojected timelines for TRL 8 was collected basedon manufacturer/developer references. These were

Area Technology TRLPropulsion High-speed turbine 6

Dual-mode scramjet 7Continuous Rotating Detonation 4Rocket-based combined cycle 3-4Turbine-based combined cycle 5-6Turbine-rocket based combined cycle 2-3Magneto-hyodrodynamic drive 3SABRE 3Scimitar 3Pre-cooled turbojet 4-5

Material Ultra-high temperature ceramics 4-5Ceramic Matrix Composites 7Metallic Matrix Composites 7Polymer nanocomposites 3-4Boron nitride nanocomposites 4Reinforced Carbon Carbon 9

TPS High-performance heat pipe 6Electron transpiration cooling 2Structurally Integrated TPS 3-4TUFROC 9Opacified Fibrous Insulation 4Internal Multi-screen Insulation 4

Aerodynamics Hypersonic I-plane aircraft configuration 1-2Distributed roughness 3

Sonic mitigation Configuration shaping 1-2Plasma boom optimization 2-3

Table 1: Novel concepts and respective TRL

corrected based on personal interpretation, takinginto account how old the references were and whatprogress has been recently achieved for each tech-nology. All this information is compiled in the finaltechnology roadmap timeline displayed in Figure 2.

3. Implementation3.1. Test case

After concluding the technology roadmap and iden-tifying high-potential technologies for hypersonicapplications, the goal is to create a numerical modelfor some of those technologies in SUAVE, a frame-work tool developed by Embraer and Stanford Uni-versity, which currently has no hypersonic capabil-ities. By doing so, it is possible to assess the im-pact of novel technologies in determining the mini-mum requirements to transport a specified payload.Since the main focus of this analysis is placed in theindividual performance of hypersonic technologies,there is a need for a vehicle-oriented assessmentrather than mission-oriented. To provide a morerobust analysis, the same vehicle is applied to twodifferent test cases, facing distinct conditions: anhypersonic cruise vehicle and an ascent-and-reentryvehicle.

A small hypersonic aircraft is envisioned, herebyreferenced as Hypersonic Multi-purpose Vehicle(HMV). Both the air-breathing cruise and theascent-and-reentry variations (hereby referenced asHMV-CAV and HMV-ARV, respectively) share theexact same design. For its initial sizing, the Lock-heed SR-72 was used as a baseline.

Despite being designed to fulfill the same goal,HMV-CAV and HMV-ARV operate in very differ-ent regimes given the clear distinction in freestream

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Figure 2: Proposed roadmap

(a) Top view (b) Isotropic perspective

Figure 3: HMV final design

conditions each vehicle faces.

For the case of the HMV-CAV, it has an appar-ently similar mission profile to that of a regularsubsonic airliner. However, the climb segment isusually done at constant dynamic pressure to avoidexcessive loads on the vehicle structure during thishigh-speed maneuver

For the HMV-ARV, the spacecraft will follow asimilar trajectory to that of the HMV-CAV duringthe initial climb phase, at constant dynamic pres-sure. Upon reaching a sufficiently high Mach num-ber and altitude, air-breathing operation is shutdown and the rocket engine kicks in; this marksthe beginning of a final ascent phase up until max-imum altitude is reached. From here, the re-entry

phase takes place.Upon reviewing the roadmap and analyzing the

predicted mission profiles, it was decided the bestengines to model for each test case would be aturbine-based combined cycle (HMV-CAV) and aturbine-rocket based combined cycle (HMV-ARV).

3.2. Numerical modelsSUAVE does not incorporate any numerical meth-ods targeting hypersonic flight in particular. Cur-rent supersonic capabilities extend to propulsionand aerodynamic correlations based on empirical orsemi-empirical laws. Therefore, after analyzing theactual SUAVE capabilities and cross-checking themwith the basic requirements to run the academictest case, a number of numerical methods were la-beled as necessary to perform a basic hypersonicanalysis.

Weight distribution Hypersonic aircrafts re-quire complementary structures that enable themto fly at such speeds (e.g., TPS, cryogenic stor-age tanks, structural reinforcements). As such,the empirical relations currently used for subsonicaircraft design might fail to correctly estimate itsweight. SUAVE has several different methods foraircraft weight estimation depending on airframedesign (i.e., tube and wing, blended wing body),even accounting for out-of-the-ordinary propulsionsystems (i.e., human and solar-powered). Becausenone of them fit the test case requirements, the Hy-personic Aerospace Sizing Analysis (HASA) weightestimation method was used [12].

LOX/LH2 rocket engine To design a low-fidelity model for preliminary rocket performance,the ideal rocket theory [9, 20] is applied, which pre-dicts that specific impulse is a function of a dragcoefficient, CD and thrust coefficient, CF .

Isp =CF

CD · g(1)

Given the complex relations between variables, asingle model for liquid rocket propulsion was cre-ated using exclusively a propellant combination ofLOX and LH2. Combustion parameters are setbased on Braeuning [6]: using the plots of adia-batic flame temperature, gas molecular weight andspecific heat ratio with respect to chamber pres-sure and O/F , polynomial and logarithmic func-tions were built using EXCEL and used for inter-polation.

To verify the numerical model, data on severalLOX/LH2 engines was collected [7, 23, 10, 11]. Thesample is wide and diverse, contemplating enginesfrom different decades purposefully built with dis-tinct on-design conditions (i.e.g, RD-0120 is a first-

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stage engine whereas Vulcain 2 is a second-stage en-gine). The next step was to use information com-piled from Braeuning to calculate specific propel-lant properties and obtain values for specific im-pulse, which were then compared to experimentaldata on each engine, both at sea level (if applica-ble) and vacuum. Figure 4 provides a visual repre-sentation of the validation process for all engines,including the parameters used in the calculations.

Figure 4: LOX/LH2 rocket engine model valida-tion, error bars of 5%

Ramjet The ramjet model was expanded uponthe existing turbojet model available in SUAVE.New constraints were added (e.g., thermal choking)and underlining simplifications were assumed. Forinstance, the flow is considered steady, quasi-one di-mensional, purely axial and behaves as perfect gas.The engine is comprised of a diffuser, burner andexit nozzle, a seen in Figure 5.

Figure 5: Ramjet engine schematics [13]

For the purpose of this demonstration, a simplediffuser design is implemented, using only one nor-mal shock to decelerate the flow. This initial ap-proach is much less efficient and results in a con-servative estimate for the diffuser performance. Us-ing Rayleigh flow equations, there is a cap on theheat transfer to the flow in the combustion chamber.This translates to a maximum exit temperature re-lated to thermal choking, TR. There are also tem-perature restrictions related to the materials useddownstream of the burner; for turbojets, this is ahigh pressure turbine, where mechanical and ther-mal stresses are very high due to rotation; for aramjet engine, this is the nozzle chamber, whichwithstands much higher temperatures. Finally, the

flow exits through the expansion nozzle, where it isconsidered to be fully expanded.

The new SUAVE ramjet model is verified withtheoretical quasi-one dimensional performance pre-dictions of specific impulse. Firstly, a range of Ispfor ramjet engines was obtained for hydrocarbonand hydrogen fuels. Then, general SUAVE efficien-cies were set, the same for both fuel types. Thesevalues ranged from 95 to 100% to move away fromthe ideal ramjet performance. To compare the nu-merical model with the expect Isp range, JP-7 andLH2 were chosen for the simulation; that is, data onthe fuel combustion heat, stoichiometric fuel-to-airratio and energy density was collected [22, 8].

Figure 6: Ramjet model verification

Scramjet To design a low-fidelity model ofscramjet engine, core assumptions were consideredto reduce the number of constraints to the problem.The Heiser and Pratt methodology [14] was applied;this implies the use of cycle static temperature ratioto model the inlet, a constant-pressure combustionwith a fixed stoichiometric fuel-to-air ratio in theburner and a fully expanded flow in the nozzle.

Figure 7: Scramjet engine schematics [19]

Since scramjet technology is still an active, high-priority research topic, there is a lack of pub-licly available experimental data to validate thismodel. There are, nonetheless, numerous scram-jet numerical models available. For the verifica-tion process, three different scramjet codes wereused: Ramjet Performance Analysis Code (RJPA),Simulated Combined-Cycle Rocket Engine Analysis

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Module (SSCREAM) and VTMODEL. The verifi-cation process used liquid hydrogen as fuel with sto-ichiometric fuel-to-air-ratio. The SUAVE model pa-rameters and the numerical model comparison canbe consulted in Figure 8.

Figure 8: Scramjet model verification

Most of the combustion-related parameters wereassumed from Heiser and Pratt [14] as they haveno equivalent to any of the codes used for com-parison. The results highlight some of the dis-crepancies that exist between other numerical mod-els and demonstrate the scramjet sensibility to mi-nor modifications in flow properties and componentefficiency. Nevertheless, the newly implementedSUAVE scramjet model provides a specific impulseestimate either within the expected range or witha small deviation from it, which is sufficient for apreliminary design basis.

Aerothermodynamic and re-entry modelThere are several aerothermodynamic models thathelp predict temperatures and heat fluxes acrossall sections of an airframe. Higher-fidelity mod-els rely on CFD software but are unnecessarilycomputational-heavy for preliminary design. Fora simpler approach, the aerothermodynamic anal-ysis in SUAVE will be performed based on semi-empirical laws. Perhaps the simplest methodfor estimating hypersonic aerodynamic heating isto use Equation 2, an approximation for trans-atmospheric vehicles derived from Anderson [16]

qw = ρ∞0.5 · V03 · 1.83 × 10−8 · rLE

−1/2 (2)

QW =

∫ t=tf

t=0

qwdt (3)

In SUAVE, mission profiles are built from dif-ferent segments, properly arranged in climb, cruise,hover or descent categories. For atmospheric opera-tions there are plenty of mission segments to choose

from but there are none available for spacecraft tra-jectories, including re-entry. For the HMV-ARV, itis necessary to create such segments to study heatflux distribution during re-entry. For maximizingthe spacecraft range, a lifting trajectory model ischosen. In a lifting reentry, there are four mainforces being exerted on the aircraft: weight, drag,lift and centripetal force. Equations for velocity andacceleration throughout re-entry are listed below:

v = vre ·

[1 +

ρoREarth

2β

L

De−h/hs

]−1

2(4)

n = − 1

2β

ρoREartheh/hs +

L

D

(5)

The IXV spacecraft was used as reference tovalidate both the atmospheric re-entry and theaerothermodynamic model and to do so requiresvehicle and trajectory information [5]. The finalresults of the validation process are displayed inFigure 9. The reentry profile matches the actualflight data (expected re-entry time of ≈ 20 min-utes) and the maximum experimental heat flux isof 65 W/cm2, which corresponds to a 4.6% relativeerror for a CD = 0.4.

Figure 9: Atmospheric and aerothermodynamicmodel validation, for CD = 0.4 (red) and CD = 0.7(green)

4. ResultsResults for the HMV-CAV and HMV-ARV are pre-sented below.

4.1. Air-breathing cruise (HMV-CAV)Mission profile selection For a hypersonic air-breathing cruise and acceleration vehicle, climb isdone at constant dynamic pressure. When this pa-rameter is fixed, the cruise altitude (influenced bythe freestream density) automatically determinesthe cruise speed. On the other hand, for hypersonicflight, the cruise Mach number must be at leasthigher than 4.0. Therefore, there is range of val-ues for dynamic pressure (between 25 and 80 KPa)

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and cruise altitude (between 20 and 25 km) thatfulfill all the requirements.

From a performance standpoint, there is a gen-eral interest in reducing the flight duration but alsomaximizing the flight range, with a fixed amountof fuel. The optimum cruise design point is there-fore chosen based on the best trade-off betweenthese two variables, represented by the utility func-tion, U , shown in Equation 6. For each coordinate(q, h), the values for the maximum and minimumrange and duration are stored and the cruise dis-tance is iteratively increased until a final fuel mar-gin of (2.50±0.30)% is reached. For each pair (q, h),U returns a value from 0 to 1, where 1 is the bestpossible result and 0 is the least favorable. Thefunction is plotted in Figure 10 for 225 coordinates.The final climb design conditions are set to 45 KPaand 25 km.

U(q, h) =1

2· X −Xmin

Xmax −Xmin+

1

2·

(1− t− tmin

tmax − tmin

)(6)

Figure 10: HMV-CAV utility function. Contourlines represent cruise Mach number, M

Baseline solution The final mission profile isdisplayed in Figure 11. The heat flux distributionat the stagnation point of the vehicle nose is demon-strated in the left graph of Figure 13. This pointin particular is chosen as it undergoes some of themost severe heating conditions. In the right graphof Figure 13, L/D is plotted over time; the red-dashed line represents the Kuchemann empirical re-lation [4] for L/Dmax in hypersonic flight. Figure12 displays the thrust at all points of the missionand the throttle value that was iterated to providesaid thrust. From this figure, throttle exceeds 100%in certain mission segments which indicates that theengine is not sized correctly; that is, it’s producinga higher thrust than what it is capable of. This callsfor a corrected solution.

Figure 11: HMV-CAV mission profile

Figure 12: HMV-CAV propulsion general parame-ters

Figure 13: HMV-CAV aerothermodynamic andaerodynamic analysis

Enhanced solution As mentioned above, whilein general the results obtained for the baseline so-lution are satisfactory, there is a misrepresentationof the engine performance expressed in the throt-tle function. There are various options to deal withthis situation (e.g., changing the mission profile it-self). However, the most fruitful and elegant solu-tion is to tweak the efficiency parameters of all thesubcomponents of the engine network as it can alsoserve as a discussion point for hypersonic propul-sion technology maturation. This combination ofnew efficiencies, with partial effort on the combus-tion process, is the minimum necessary to obtain avalid flight for this mission profile in particular, andrepresents the progress required for engine subsys-tems to allow that to happen. While a few modifica-tions were made to the inlet and nozzle parameters,the main performance gains were attributed to thecombustion process, in an effort to simulate futurecombustion technology enhancement, a major area

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of development in hypersonic propulsion.

Figure 14: Comparison of the optimized solution(red) and the baseline solution (blue)

Improved efficiency leads to less fuel spent, whichin turn reflects in a higher range assuming the samefinal fuel margin. The simulation is run again anda final range of 11,000 km is achieved.

4.2. Ascent and re-entry vehicle (HMV-ARV)Mission profile selection For a generic hyper-sonic ascent and re-entry vehicle, there is a higherinterest in maximizing the flight range given thatthe ascent phase always develops over a small timeframe. In this fashion, for fixed aircraft properties(e.g., area, L/D, CD), range may be significantlyincreased by achieving a higher altitude and/orspeed at the end of the ascent phase since most ofthe flight range develops over the re-entry process.Therefore, through the built-in SUAVE optimiza-tion functions, it is possible to plot the range for aninput array of initial re-entry altitude and speed,(h, v). For the purpose of this test case, the re-entrysegment takes place immediately after the comple-tion of the orbital ascent segment and, therefore,these input conditions are simultaneously the finalclimb altitude and speed, respectively.

To simulate the re-entry segment, estimates forL/D and CD have to be provided. These were ei-ther based on the IXV or the Space Shuttle orbiterreferences [21].

Unlike the HMV-CAV cruise design trade-off,there is no point in using a similar utility functionbecause the ascent to orbit develops over a rela-tively small period of time when compared to thetotal flight duration; on the other hand, any ad-ditional seconds during the final ascent may leadto considerable changes to the final aircraft range,as seen in Figure 15. Therefore, only the range isused as a selection criteria. The climb parametersused for the HMV-ARV are the same that have beenselected for the HMV-CAV, including constant dy-namic pressure for the air-breathing climb segment.The final climb design point has a final orbital speedof 6.0 km/s at an altitude of 100 km.

Baseline solution The final mission profile isrepresented in Figure 16. Most of the climb seg-

Figure 15: Range and duration for the HMV-ARV.Contour lines represent fuel margin

ments are the same as that of the HMV-CAV ex-cept for the final ascent to orbit; since no gravityturn segment exists in SUAVE, this is replaced bya linear climb at linear Mach numbers.

Weight distribution is shown in Figure 17. Forthe most part, fuel consumption is moderate, inline with the HMV-CAV case, until rocket opera-tion kicks in; now the propellant consumption israpidly increasing as both on-board oxidizer andfuel are being used. The main concern from theHMV-CAV base solution is replicated: the engineis poorly sized for the mission profile and throttleovershoots by 50%. This calls for an enhanced so-lution.

Figure 16: HMV-ARV mission profile

Figure 17: HMV-ARV weight distribution

Enhanced solution For the enhanced solution,the same logic from the HMV-CAV optimizationcase applies; that is, the burner parameters wereimproved to obtain a more efficient combustion. Inthis way, it may be possible to properly size theengine and correct the throttle issue from the pre-

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Figure 18: HMV-ARV propulsion general parame-ters

vious subsection. Through fuel savings it may alsobe possible to reach higher speeds and significantlyincrease the spacecraft range. For the common air-breathing cycle, the improvements are exactly thesame presented for the HMV-CAV test case; rocketparameters remain unaltered. Rockets are extraor-dinarily sensible to weight and therefore any fuelsavings are predicted to have a significant impactin the spacecraft performance. Upon improving theindividual component efficiencies, the new climb de-sign conditions are set for a final climb speed of 7.5km/s at an altitude of 100 km.

Figure 19: HMV-ARV optimized mission profile

Figure 20: HMV-ARV optimized aerothermody-namic and aerodynamic distribution

5. ConclusionsThe technology roadmap proved most useful inidentifying and assessing various technological solu-tions in a multitude of technological areas: propul-sion, aerodynamics, materials, thermal protectionsystems and control. To better understand if andhow these new technologies may be incorporatedin the future, a brief explanation for each one was

Figure 21: HMV-ARV optimized propulsion specificparameters

Figure 22: HMV-ARV optimized propulsion generalparameters

presented and used to discuss their application in avariety of hypersonic aircraft.

Regarding the SUAVE simulations, for the samefuel margin, any of the HMV-CAV simulations out-class their respective HMV-ARV counterpart whenlooking at range over duration; the range for theHMV-ARV is slightly shorter but time savings cango up as much as half the airbreathing cruise sce-nario. This is consistent with SpaceX’s predic-tions for sub-orbital flight when applied to Earth-to-Earth transport, which is one of the first applica-tions for its future BFR rocket [18]. However, oneof the many issues of this concept is not only thesafety concerns and lack of legislation for commer-cial activities, but also the extreme aerothermody-namic environment these vehicles endure. The heatload is considerably shorter but that is mainly be-cause travel time has been almost cut down in half.Heat fluxes at the nose stagnation point can go ashigh as 173% that of the HMV-CAV, due to the ex-treme speeds during the final ascent. This corrob-orates the need for high-performance materials andTPS for the ARV scenario and justifies the currentfocus on UHTC and other technologies presentedthroughout the roadmap development.

SUAVE is now able to run a very preliminaryanalysis, which can improved upon by updating thenew models that have just been introduced. For in-stance, the scramjet model may be improved to in-clude an isolator to simulate shock trains and chem-ical equilibrium equations may also be added tomore accurately predict the combustion tempera-tures. On the other hand, the rocket model canbe further expanded to include other propellant

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combinations, using the Braeuning references. Theaerothermodynamic model can also be be improvedto give out more accurate predictions of heat distri-bution.

Acknowledgements

The author would like to acknowledge the helpprovided by the SUAVE Development Team overthe SUAVE forums for their helpful insight on theSUAVE methodology and assistance with multiplePython and algorithm-related issues. The authorwould also like to acknowledge the support of EM-BRAER for providing an extensive bibliographyand guidance throughout the development of thispaper.

IPFN activities received financial support fromFundacao para a Ciencia e Tecnologia throughproject UID/FIS/50010/2013

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