Aero-thermodynamic design of JAXA’s hypersonic passenger aircraft

Atsushi UENO and Hideyuki TAGUCHI Japan Aerospace Exploration Agency 1st International Symposium: “Hypersonic flight: from 100.000 to 400.000 ft” Rome, Italy 30 June, 2014 1

Contents

1.

Hypersonic research at JAXA   R&D roadmap Baseline configuration defined by MDO 2.

Aerothermodynamic design  Evaluation of aerodynamic heating  Comparison between CFD and WTT  Brief introduction of TPS design 3.

Hikari project (Europe-Japan Collaboration)  Brief introduction of Hikari’s results  Evaluation of hypersonic engine performance 4.

Summary 2

1. Hypersonic research at JAXA

 Hypersonic research at JAXA Balloon-based Operation Vehicle 3 Hypersonic Integrated Control Experiment HIMICO Hypersonic Technology Experiment HYTEX Hypersonic Business Jet TSTO Hypersonic Transport

Mach 2 Mach 5 Mach 5 Mach 0 ~ 5

Small PCTJ (Mach2) Small PCTJ (Mach 5) Medium PCTJ Large PCTJ

JAXA’s R&D Roadmap on Hypersonic Transport Aircraft

Variable intake Pre-cooler Core engine

Pre-Cooled TurboJet Engine (PCTJ)

Variable nozzle

4500 4000 3500 3000 2500 2000 1500 0 68 66 64 62 60 58 56 54 0 1400 a 1200 1000 800 600 400 6.5

6 5.5

5 4.5

4 3.5

3 2000 4000 時間[sec] 6000 8000 200 0 2000 6000 8000 2.5

0 4000 時間[sec]

1. Hypersonic research at JAXA

2000  x 10 -6 1 Hypersonic transport – 100 passengers 重心 0.5

– Mach 5 / Altitude 25 km 30 25 Engine cut-off 4000 時間[sec] 20 6000 2000 4000 時間[sec] 6000 8000 -1.5

-2 -2.5

-3 0 15 10 5 2000 3. Cruise (around Mach 5) 4000 6000 8000 時間[sec] 0 0 1 8000 4 Const. dynamic pressure 2 3 4 5 6 2. Acceleration 4. Deceleration (90 min., 7600km) 1. Take-off (10 min., 400km) (10 min., 700km) 5. Landing Pacific ocean

Mission profile

1. Hypersonic research at JAXA

 Baseline configuration – Multidisciplinary design optimization

Shape Weight

Shape

Aero.

AoA Mach Weight Inlet area Aero. force

Propulsion

Thrust SFC Altitude Mach

Mission

Fuel weight

Baseline configuration

Design variables

Optimization

Objective function Constraint function MTOW

Baseline specifications

370 ton Dry Weight Fuel Weight Length 190 ton 180 ton 87 m Span Wing Area Engine Thrust (SLS) 35 m 770 m 2 PCTJ 44 ton X 4 5

2. Aero-thermodynamic design

 Evaluation of Aerodynamic heating rate – In MDO, aero. heating was not taken into account.

– TPS weight was estimated using empirical relation.

• HASA, NASA-Contractor Report 182226  CFD and wind tunnel test (WTT) were conducted to evaluate aero. heating.

CFD (WTT condition) WTT 6 (Flight condition) TPS design

2. Aero-thermodynamic design

 CFD analysis – Navier-Stokes analysis • JAXA’s UPACS code – Equation: RANS – Flux discretization: AUSMDV (3rd order) – Turbulent model: Spalart-Allmaras – Number of points: 15 million • Flow condition: – Wind tunnel condition » T0 = 700 [K], M = 5, AoA = 5 [deg] » Re = 1.7x10

6 (P0=1.0 [MPa]), Laminar » Re = 7.1x10

6 (P0=1.5 [MPa]), Turbulent » Tw = 303 [K], Isothermal wall – Flight condition » h = 24.2 [km], M = 5, AoA = 5 [deg] » Re = 4.0x10

8 , Turbulent » Tw = 823 [K], Isothermal wall Validation TPS design 7

2. Aero-thermodynamic design

 Wind tunnel test – JAXA HWT1 Type Test section Mach number Nozzle exit Max. duration

HWT1 HWT2

Blow down / vacuum intermittent Free jet 5, 7, 9 φ0.5m

120sec 10 φ1.27m

60sec 8

2. Aero-thermodynamic design

 Wind tunnel test – Wind tunnel model Material

0.25% model 0.74% fuselage model

Vespel (polyimide plastic) M, AoA P0, T0 Re Measurement M = 5, AoA = 5 [deg] 1.0 [Mpa], 700 [K] 1.7x10

6 , Laminar 1.5 [MPa], 700 [K] 7.1x10

6 , Turbulent Temperature (IR thermography) 9

Temperature

semi-infinite, 1D heat equation

Aerodynamic heating

sphere ( Φ1mm) Boundary layer trip 0.25% model, L=220mm Fuselage + Wing + V-tail (q w on all components in laminar B.L.) 0.74% model, L=643mm Fuselage (q w in turbulent B.L.)

Wind tunnel model

2. Aero-thermodynamic design

 Result of WTT – Result of 0.25% model ( Laminar boundary layer ) • Wind tunnel test 10

Aerodynamic heating (M = 5, AoA = 5 [deg], Upper surface) Aerodynamic heating on all components was measured.

Large aerodynamic heating due to separated vortex was observed.

2. Aero-thermodynamic design

 Comparison between CFD and WTT – Result of 0.25% model ( Laminar boundary layer )

Upper surface (WTT)

11

Upper surface (CFD)

q w upper Nose

Lower surface (WTT)

q w lower semi-infinite,1D heat equation is not correct.

 Overestimation in WTT

Lower surface (CFD) Distribution of Stanton number at AoA=5deg.

CFD agrees with wind tunnel test qualitatively except in region where thickness of model is thin.

2. Aero-thermodynamic design

 Comparison between CFD and WTT – Result of 0.74% fuselage model ( Turbulent boundary layer )

Upper surface (WTT)

Boundary layer trip Camera #1 Camera #2

Upper surface (CFD) Distribution of Stanton number at AoA=5deg.

Boundary layer transition was observed behind boundary layer trip.

High aero. heating due to separated vortex was observed also in turbulent B.L.

12

2. Aero-thermodynamic design

 Comparison between CFD and WTT – Result of 0.74% fuselage model ( Turbulent boundary layer ) 13 St 0.0010

0.0008

0.0006

0.0004

0.0002

Boundary layer trip CFD WTT 0.2

0.0

St 0.0010

0.0008

0.0006

0.0004

0.0002

0.0

0.2

CFD WTT 0.4

Camera #1

Center of fuselage

x/L 0.6

0.4

x/L 0.6

Camera #2 0.8

Center of vortex (y/L=0.028)

0.8

1.0

1.0

Aero. heating differs in the region where separated vortex is attached.

CFD shows larger aero. heating.

2. Aero-thermodynamic design

 TPS design based on CFD result High heating rate (q w : ~ 100kW/m 2 ) Upper 14 High heating rate (q w : ~ 100kW/m 2 ) Lower Cryogenic tank (q w : 5 ~ 20kW/m 2 ) Cabin (q w : 5 ~ 15kW/m 2 ) Thin wing (q w : 5 ~ 30kW/m 2 )

CFD result at flight condition

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/m 2 .

 Summary  Results 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.

Objectives: Task of JAXA: Status:

3. Hikari project

Europe-Japan “HIKARI” Collaboration

Market analysis, Environmental Impact Assessment, Aircraft Systems Study, Propulsion, Common R&D Roadmap Performance evaluation of Hypersonic Pre-Cooled Turbojet Engine Mach 4 experiment has been successfully conducted.

Performance map will be provided to research partners in August.

15 Mach 4 Direct Connect Test -High Temperature Structure -Mach 4 Operation 2005 Mach 4 Wind Tunnel Test 2015 -Starting Sequence -Heat Structure of Variable Mechanism 2020 2025

4. Summary

 Hypersonic 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.

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