PowerPoint

Aero-thermodynamic design of JAXAs hypersonic passenger aircraft1st International Symposium: Hypersonic flight: from 100.000 to 400.000 ftRome, Italy30 June, 2014

Atsushi UENO and Hideyuki TAGUCHIJapan Aerospace Exploration Agency#Thank you, chairman. Id like to introduce research activity on hypersonic transport at JAXA.1ContentsHypersonic research at JAXAR&D roadmapBaseline configuration defined by MDO

Aerothermodynamic designEvaluation of aerodynamic heatingComparison between CFD and WTTBrief introduction of TPS design

Hikari project (Europe-Japan Collaboration)Brief introduction of Hikaris resultsEvaluation of hypersonic engine performance

Summary#In this presentation, I will focus on aero-thermodynamic design. We have performed CFD analysis and wind tunnel test to evaluate aerodynamic heating, and thermal protection system was designed. In addition, Hikari project (Europe-Japan collaboration) is briefly introduced.21. Hypersonic research at JAXAHypersonic research at JAXAMedium PCTJ

Large PCTJMach 0 ~ 5Mach 5Mach 5Mach 2

Small PCTJ (Mach 5)Small PCTJ (Mach2)

TSTOHypersonic Integrated Control ExperimentHIMICOHypersonic Technology ExperimentHYTEXBalloon-based Operation Vehicle

Hypersonic Business JetHypersonic Transport

Variable intakePre-coolerCore engine Variable nozzlePre-Cooled TurboJet Engine (PCTJ)JAXAs R&D Roadmap on Hypersonic Transport Aircraft#This figure shows JAXAs R&D roadmap on hypersonic transport. We have developed hypersonic turbojet engine named PCTJ, pre-cooled turbojet engine. Owing to this pre-cooler, this engine can be operated at up to Mach 5. Currently we are developing small-sized PCTJ. Length of engine is about 2m. Using this engine, we plan to conduct flight test at Mach 5. After that, we are aiming at hypersonic business jet using medium-sized engine. Then we are aiming at large-sized aircraft. I will introduce large-sized aircraft.

Memo. HIMICO: Rocket-launched (JAXAs sounding rocket), L=1m, ram jet, control of inlet, flight controlHYTEX: Rocket-launched (JAXAs epsilon rocket), L=6m, small PCTJ, API, engine performance31. Hypersonic research at JAXAHypersonic transport100 passengersMach 5 / Altitude 25 km2 hours from Tokyo to Los AngelesUse existing airports2. Acceleration3. Cruise (around Mach 5)1. Take-off

Pacific ocean

4. Deceleration5. Landing(10 min., 400km)(90 min., 7600km)(10 min., 700km)

Mission profile

MachAltitude [km]Engine cut-offConst. dynamic pressure#The number of passenger is 100. Cruise Mach number is 5 at an altitude of 25km. We are aiming at trans-pacific flight from Tokyo to Los Angeles. This figure shows typical flight trajectory. Take-off from existing airport. Ascent along flight path of constant dynamic pressure. Hypersonic cruise. During descent, engine is cut-off. Finally landing. It will take about 2 hours.41. Hypersonic research at JAXABaseline configurationMultidisciplinary design optimizationWeightBaseline specificationsMTOW370 tonDry Weight190 tonFuel Weight180 tonLength87 mSpan35 mWing Area770 m2EnginePCTJThrust (SLS)44 ton X 4Baseline configurationInlet areaAero. forceThrustSFCAltitudeMachAoAMachShapeAero.ShapeWeightPropulsionMissionFuel weightOptimizationDesign variablesObjective functionConstraint function

Fuel tankFuel tankCabin#Baseline configuration of large-sized aircraft was defined by using MDO technique. The aerodynamic shape and engine size were optimized to satisfy mission requirements such as range performance and flight time. Optimized length is about 90m. Weight is about 400 metric ton. We have two fuel tanks at front and rear part of fuselage in order to control the position of center of gravity. The cabin is located at center of fuselage. For this baseline configuration, detailed aero-thermodynamic design was conducted.52. Aero-thermodynamic designEvaluation of Aerodynamic heating rateIn MDO, aero. heating was not taken into account.TPS weight was estimated using empirical relation.HASA, NASA-Contractor Report 182226CFD and wind tunnel test (WTT) were conducted to evaluate aero. heating.CFD(WTT condition)

(Flight condition)WTTValidationTPS designAero. heating

#In the MDO, aerodynamic heating was not taken into account. TPS weight was estimated using empirical relation. Here, CFD and wind tunnel test were conducted to evaluate aerodynamic heating. In the wind tunnel test, it is difficult to realize flight condition, because Reynolds number at flight condition is too large. So at first, CFD was validated by wind tunnel test, and then, based on the CFD result at flight condition, TPS was designed.62. Aero-thermodynamic designCFD analysisNavier-Stokes analysisJAXAs UPACS codeEquation: RANSFlux discretization: AUSMDV (3rd order)Turbulent model: Spalart-AllmarasNumber of points: 15 million

Flow condition:Wind tunnel conditionT0 = 700 [K], M = 5, AoA = 5 [deg]Re = 1.7x106 (P0=1.0 [MPa]), LaminarRe = 7.1x106 (P0=1.5 [MPa]), TurbulentTw = 303 [K], Isothermal wall

Flight conditionh = 24.2 [km], M = 5, AoA = 5 [deg]Re = 4.0x108, TurbulentTw = 823 [K], Isothermal wallTPS designValidation

#Method of CFD analysis is shown here. JAXAs UPACS code was used. Firstly, CFD was conducted at wind tunnel condition. Wall condition was isothermal wall and wall temperature was set at about 300K considering the wall temperature of wind tunnel model. This result was compared to that of wind tunnel test as a validation. Then, CFD was conducted at flight condition. The wall temperature was set at about 800K considering the usable temperature of TPS material. Based on this result, TPS was designed.72. Aero-thermodynamic designWind tunnel testJAXA HWT1

HWT1HWT2TypeBlow down / vacuum intermittentTest sectionFree jetMach number5, 7, 910Nozzle exit0.5m1.27mMax. duration120sec60sec#Wind tunnel test was conducted at JAXAs hypersonic wind tunnel. Free stream Mach number is 5. Diameter of test section is 0.5m. The maximum duration is 120 seconds.82. Aero-thermodynamic designWind tunnel testWind tunnel model

Wind tunnel model

Boundary layer trip0.25% model0.74% fuselage modelMaterialVespel (polyimide plastic)M, AoAM = 5, AoA = 5 [deg]P0, T01.0 [Mpa], 700 [K]1.5 [MPa], 700 [K]Re1.7x106, Laminar7.1x106, TurbulentMeasurementTemperature (IR thermography)sphere (1mm)TemperatureAerodynamic heatingsemi-infinite,1D heat equation0.25% model, L=220mmFuselage + Wing + V-tail (qw on all components in laminar B.L.)0.74% model, L=643mmFuselage (qw in turbulent B.L.)#We have two wind tunnel models. This smaller model is a full configuration model having fuselage, wing, and vertical tail. Reynolds number was less than transition Reynolds number, and boundary layer was laminar. However, at flight condition, boundary layer becomes turbulent. So we have to measure the aerodynamic heating for turbulent boundary layer. In order to increase the Reynolds number, in this model, length was enlarged to about 600mm. Only the fuselage was modeled. P0, stagnation pressure, was increased from 1.0 to 1.5 MPa. As a result, the Reynolds number becomes larger than transition Reynolds number. In addition, boundary layer trip was attached near the nose. In this test, temperature on the model surface was measured by using Infra-red thermography. Then, assuming the semi-infinite 1D heat equation, the aerodynamic heating was evaluated.

92. Aero-thermodynamic designResult of WTTResult of 0.25% model (Laminar boundary layer)Wind tunnel testAerodynamic heating on all components was measured.Large aerodynamic heating due to separated vortex was observed.Aerodynamic heating(M = 5, AoA = 5 [deg], Upper surface)

#Result of wind tunnel test for smaller model is shown here. Boundary layer is laminar. This figure shows aerodynamic heating on the upper surface of the model. Large aerodynamic heating can be seen here. This is due to the vortex. Flow is separated at side edge of the fuselage and vortex is generated. This vortex thins the boundary layer thickness. As a result, aerodynamic heating is increased in the region where this vortex is attached.102. Aero-thermodynamic designComparison between CFD and WTTResult of 0.25% model (Laminar boundary layer)

Upper surface (WTT)Upper surface (CFD)CFD agrees with wind tunnel test qualitatively except in region where thickness of model is thin.Lower surface (WTT)Lower surface (CFD)Distribution of Stanton number at AoA=5deg.qwqwupperlowerNosesemi-infinite,1D heat equation is not correct.Overestimation in WTT

#This figure shows the comparison between CFD and wind tunnel test. They agree qualitatively. However, in the wind tunnel test, aerodynamic heating is slightly overestimated around the nose and wing leading edge. This overestimation is due to the thickness of the wind tunnel model. When the thickness is small, it is difficult to distinguish between heating on upper surface and heating on lower surface. So this assumption is not correct. As a result, aerodynamic heating is overestimated in the region where the thickness of wind tunnel model is small. But CFD and wind tunnel test showed qualitative agreement.112. Aero-thermodynamic designComparison between CFD and WTTResult of 0.74% fuselage model (Turbulent boundary layer)

Camera #1Camera #2

Upper surface (WTT)Upper surface (CFD)Distribution of Stanton number at AoA=5deg.Boundary layer tripBoundary layer transition was observed behind boundary layer trip.High aero. heating due to separated vortex was observed also in turbulent B.L.#This figure shows result of larger model having only the fuselage. In the downstream of boundary layer trip, boundary layer becomes turbulent and aerodynamic heating is increased. In this region, effect of vortex can be seen also in the turbulent boundary layer. Wind tunnel test and CFD agree qualitatively. To compare more precisely, 2 dimensional plots are shown.-----------------------------------------------From this figure, in the wind tunnel test, the vortex seems to be diffused more widely. We think difference in the diffusion of vortex may affect the aerodynamic heating. 12

2. Aero-thermodynamic designComparison between CFD and WTTResult of 0.74% fuselage model (Turbulent boundary layer)

Center of fuselageCenter of vortex(y/L=0.028)CFDWTTCFDWTTBoundary layer tripCamera #1Camera #2Aero. heating differs in the region where separated vortex is attached.

CFD shows larger aero. heating.#This is the aerodynamic heating along x-axis. Along center line and along this line. Here, effect of vortex can be seen. In this region, boundary layer is laminar in wind tunnel test and aerodynamic heating is lower than that of CFD analysis. In the downstream of the boundary layer trip, boundary layer becomes turbulent and aerodynamic heating shows good agreement with that of CFD analysis. On the other hand, along this line, CFD analysis shows larger aerodynamic heating. We think this difference is due to vortex. ------------------------------------------------------------------As seen here, CFD analysis shows larger aerodynamic heating. So the TPS was designed based on the result of CFD analysis to be on the safe side.

132. Aero-thermodynamic designTPS design based on CFD result

UpperLower

High heating rate(qw: ~ 100kW/m2)Cryogenic tank(qw: 5 ~ 20kW/m2)Cabin(qw: 5 ~ 15kW/m2)Thin wing(qw: 5 ~ 30kW/m2)High heating rate(qw: ~ 100kW/m2)Super alloy (Inconel) honeycomb should be applied in the region where aerodynamic heating is large (e.g., nose and leading edge).

Ti multi-wall can be applied in the region where qw is about 20kW/m2.SummaryResults of wind tunnel test and CFD agreed qualitatively.CFD showed larger aerodynamic heating in the region where separated vortex is attached.Different turbulent model should be tested in the future.TPS was designed based on aerodynamic heating obtained by CFD.TPS material was selected.

CFD result at flight condition#Based on this CFD result, TPS was designed using 1 dimensional analysis. The method is not shown here, I will introduce only the results. In the TPS design, material of TPS was selected and thickness of TPS was determined. At nose and wing leading edge, aerodynamic heating is nearly 100kw/m2. In these regions, super alloy (Inconel) honeycomb should be applied. Other than these regions, aerodynamic heating is about 20kw/m2. In these regions, ti multi-wall can be applied. Here is the summary:143. Hikari project Europe-Japan HIKARI Collaboration

20102025202020052015

Mach 4 Direct Connect Test

High Temperature StructureMach 4 OperationMach 4 Wind Tunnel Test

Starting SequenceHeat Structure of Variable Mechanism

Hypersonic Pre-Cooled Turbojet Engine (JAXA)Objectives: Market analysis, Environmental Impact Assessment, Aircraft Systems Study, Propulsion, Common R&D RoadmapTask of JAXA:Performance evaluation of Hypersonic Pre-Cooled Turbojet EngineStatus:Mach 4 experiment has been successfully conducted.Performance map will be provided to research partners in August.#Hikari project is briefly introduced. On Japanese side, industry association, engine company, university, and JAXA join this collaboration. We have several objectives. Among these objectives, JAXA is engaged in propulsion design. Our task is to evaluate performance of hypersonic PCTJ engine. This picture shows Mach 4 direct connect test, and this year Mach 4 wind tunnel test was successfully conducted. Performance map will be provided to research partners in August.154. SummaryHypersonic passenger aircraft was studied using MDO technique.Baseline was defined.

Aerodynamic heating rate was evaluated by both CFD and WTT.CFD and WTT showed qualitative agreement.TPS was designed based on aerodynamic heating rate obtained by CFD.

Results from Hikari project was briefly introduced.

Future works:Plan for experimental vehicle with small PCTJ flying at Mach 5.

#Summary is shown here. We are now planning the flight test using small PCTJ. Budget is not secured yet. But we are trying to conduct flight test in the future.16