EFD/CFD Activities in Research for Reusable Launch Vehicles

(1)

Space Transportation Systems Engineering Laboratory, Kyushu University

EFD/CFD Activities in Research for Reusable Launch Vehicles

ౣ૶↪ဳቝቮᓔㆶᯏߦะߌߚ⎇ⓥߦ߅ߌࠆEFD/CFDߩขࠅ⚵ߺ

Yasuhiro TANI, Shigeru ASO, Kentaro HAYASHI, Kei INOUE, Kohei YAMAGUCHI and Takuro ISHIDA

⼱ ᵏኡޔ㤗↢ ⨃ޔᨋஜᄥ㇢ޔ੗਄ ᘮޔጊญ⠹ᐔޔ⍹↰ᜏ㇢

Department of Aeronautics and Astronautics, Kyushu University Fukuoka, Japan

Second Workshop of Integration of EFD and CFD, 23-24 Feb. 2009

1. Research Area in our Laboratory

2. Experimental Facilities and CFD Tools

3. Aerodynamic characteristics of RLV configuration 4. Aerodynamic heating reduction research

5. Fuel/Air mixing of SCRAM jet engine combustor

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Research Area in our laboratory

䊶High and Low Speed Aerodynamic Characteristics of Reusable Launch Vehicle (RLV)

䊶Reduction of aerodynamic heating during reentry phase for the RLVs

䊶Future aerospace propulsion system (SCRAM-jet engin, PDE )

䊶Ecological aircraft for future commuter air transportation

Research activities on RLVs

䊶Aerodynamic characteristics

fuselage cross sectional configuration 䊶Reduction of Aerodynamic heating

opposing jet, Film cooling 䊶Engines for hypersonic flight

scramjet engine, Pulse detonation engine

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Experimental Facilities in use

Test facilities in Department of Aeronautics and

Astronautics, Kyushu univ.

䊶Low Noise Low speed wind tunnel 䊶Supersonic wind tunnel

䊶Transonic wind tunnel

Test facilities in Space Transportation Systems Lab.

䊶Detonation driven Expansion tube 䊶Free piston shock tunnel

䊶Shock Tube

Other test facilities in use

䊶ISAS/JAXA Supersonic wind tunnel and Transonic wind tunnel

Former Wind tunnels in Hakozaki Campus

2m Low speed wind tunnel

15cm Supersonic wind tunnel (Mach 4)

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Kyushu Univ. Low Noise Wind Tunnel

Low Noise Test Section (No.1)

Göttingen type, in anechoic chamber 2 m width octagonal, 5 m length Max 60 m/s, 65 dB at 40 m/s L㪸㫉㪾㪼㩷Closed Test Section (No.2)

3.5m x 3.5 m (max 15m/s), 30m/s with 3.5m x 1.5m insert

Kyushu Univ. Transonic Wind Tunnel

Blow down type Transonic wind tunnel Mach 0.3䌾1.3

150mm x 450mm closed test section with slit walls

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Kyushu Univ. Supersonic Wind Tunnel

Blow down type Supersonic wind tunnel 䊶Mach 2.5 and 3.5

䊶250x200mm closed test section

Kyushu Univ. Transonic Wind Tunnel

High enthalpy Flow test apparatus 䊶Free piston shock tunnel

䊶Detonation driven Expansion tube

䊶Normal Shock Tube

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Free piston shock tunnel

High pressure tube : L 0.7 m, D 136.6 mm Compression tube : L 3.0 m, D 70 mm Piston : Mass 1.13 kg, L 49 mm Shock tube : L 3.3 m, D 60 mm

Nozzle : Exit Diameter 270 mm,Area ratio 190, Design Mach 8

Test section : Volume 1m³

CFD Tools

2D / 3D Navier-Stokes code (in house)

䊶Compressible Full Navier-Stokes / Euler 䊶Structured grid with Multi-Block formulation 䊶AUSM-DV scheme, LU-ADI

䊶Turbulence models

䊶Wilcox k-㱥 2eq. model 䊶Spalart-Allmaras 1eq. model 䊶Baldwin-Lomax algebraic model 䊶Chemical reaction

Grid Generation

䊶Gridgen

䊶Transfinite/Elliptic grid generator (In house)

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CFD Tools

Design tool

䊶PANAIR 䊶XFOIL

䊶Vortex lattice code 䊶Newtonian Flow code 䊶DATCOM

Background (1)

䊶㪭㪠㪩㪞㪠㪥㩷㪞㪘㪣㪘㪚㪫㪠㪚

㪪㫇㪸㪺㪼㩷㪪㪿㫀㫇㩷㪫㫎㫆㩷㪸㫅㪻㩷㪮㪿㫀㫋㪼㩷㪢㫅㫀㪾㪿㫋㩷㪫㫎㫆㩷䊶㪩㪦㪚㪢㪜㪫㩷㪧㪣㪘㪥㪜㩷㪲㪞㪣㪦㪙㪘㪣㪴㩷

㪩㫆㪺㫂㪼㫋㩷㪧㫃㪸㫅㪼㩷㪯㪧

㪥㪼㫎㩷㪫㪼㪺㪿㫅㫆㫃㫆㪾㫐㩷㪸㫅㪻㩷㪛㪼㫍㪼㫃㫆㫇㫄㪼㫅㫋㩷㪸㫉㪼㩷㫉㪼㫈㫌㫀㫉㪼㪻㩷㪽㫆㫉㩷㪪㫇㪸㪺㪼㩷㪫㫉㪸㫅㫊㫇㫆㫉㫋㪸㫋㫀㫆㫅㩷㪪㫐㫊㫋㪼㫄

㵬㪜㫏㫇㪼㫅㪻㪸㪹㫃㪼㩷㪩㫆㪺㫂㪼㫋㫊㵭㩽㩷㵬㪪㫇㪸㪺㪼㩷㪪㪿㫌㫋㫋㫃㪼㩷㩿㫇㪸㫉㫋㫃㫐㩷㫉㪼㫌㫊㪼㪀㵭㪚㫌㫉㫉㪼㫅㫋㩷㪪㫐㫊㫋㪼㫄㫊

㪟㫀㪾㪿㩷㪜㪺㫆㫅㫆㫄㫀㪺㩷㪜㪽㪽㫀㪺㫀㪼㫅㪺㫐㪞㫃㫆㪹㪸㫃㩷㪜㫅㫍㫀㫉㫆㫅㫄㪼㫅㫋

㪩㪼㫌㫊㪸㪹㫃㪼㩷㪣㪸㫌㫅㪺㪿㩷㪭㪼㪿㫀㪺㫃㪼䋨㪩㪣㪭䋩

http://www.rocketplane.com/

㪪㫇㪸㪺㪼㩷㫋㫆㫌㫉㫀㫊㫄㩷㫄㪸㫉㫂㪼㫋㩷㪿㪸㫊㩷㪹㪼㪼㫅㩷㫄㫆㫉㪼㩷㪸㪺㫋㫀㫍㪼㪅

http://www.virgingalactic.com/

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Background (2)

㪚㫆㫅㫊㫀㪻㪼㫉㫀㫅㪾㩷㫋㪿㪼㩷㪸㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪿㪼㪸㫋㫀㫅㪾㪸㫅㪻㩷㫊㫋㫉㫌㪺㫋㫌㫉㪸㫃㩷㫎㪼㫀㪾㪿㫋

䊶㪮㫆㫉㫊㪼㩷㪸㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪺㪿㪸㫉㪸㪺㫋㪼㫉㫀㫊㫋㫀㪺㫊㩷㫀㫅㩷㫊㫌㪹㫊㫆㫅㫀㪺㩷㫉㪼㪾㫀㫆㫅

䊶㪝㫌㫊㪼㫃㪸㪾㪼㫊㩷㪿㪸㫍㪼㩷㫄㫆㫉㪼㩷㪼㪽㪽㪼㪺㫋㫊㩷㫆㫅㩷㪸㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㫇㪼㫉㪽㫆㫉㫄㪸㫅㪺㪼㩷㪩㪣㪭㩾㫊㪮㫀㫅㪾㫊㩷㪸㫉㪼㩷㫊㫄㪸㫃㫃㩷㪺㫆㫄㫇㪸㫉㪼㪻㩷㫎㫀㫋㪿㩷㪽㫌㫊㪼㫃㪸㪾㪼㫊㩷㫉㪼㫃㪸㫋㫀㫍㪼㫃㫐

䊶㪟㫀㪾㪿㩷㩷㩷㩿㪣㪆㪛㪀 ^㫄㪸㫏 ^䊶䊶䊶

㪣㫆㫎㩷㫃㪸㫅㪻㫀㫅㪾㩷㫊㫇㪼㪼㪻㪃㩷㫉㪸㫋㪼㩷㫆㪽㩷㪻㪼㫊㪺㪼㫅㫋㪪㪿㪸㫃㫃㫆㫎㩷㪽㫃㫀㪾㪿㫋㩷㫇㪸㫋㪿㩷㪸㫅㪾㫃㪼

㩿㪠㫅㪺㫉㪼㫄㪼㫅㫋㩷㫆㪽㩷㪻㫆㫎㫅㫉㪸㫅㪾㪼㩷㩽㩷㪺㫉㫆㫊㫊㩷㫉㪸㫅㪾㪼㪀㪣㫆㫎㩷㫉㪸㫋㪼㩷㫆㪽㩷㪻㪼㫊㪺㪼㫅㫋

䊶㪟㫀㪾㪿㩷㩷㩷㪚 ^㪣 ^㫄㪸㫏 ^䊶䊶䊶

䊶㪤㪸㫅㪼㫌㫍㪼㫉㪸㪹㫀㫃㫀㫋㫐㪃㩷㪪㪸㪽㪼㫋㫐㪃㩷㩽㩷㪩㪼㫃㫀㪸㪹㫀㫃㫀㫋㫐

㪩㪼㫈㫌㫀㫉㪼㪻㩷㪸㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪺㪿㪸㫉㪸㪺㫋㪼㫉㫀㫊㫋㫀㪺㫊㩷㩽㩷㫇㪼㫉㪽㫆㫉㫄㪸㫅㪺㪼㩷㪽㫆㫉㩷㪩㪼㫈㫌㫀㫉㪼㪻㩷㪸㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪺㪿㪸㫉㪸㪺㫋㪼㫉㫀㫊㫋㫀㪺㫊㩷㩽㩷㫇㪼㫉㪽㫆㫉㫄㪸㫅㪺㪼㩷㪽㫆㫉㩷㪩㪣㪭㪩㪣㪭

Simplified fuselage models

Wind tunnel test models

䊶simplified fuselage model

䊶three different cross section

(9)

ISAS/JAXA Wind Tunnels

Transonic : M

_п

= 0.3 㨪 1.3 Supersonic : M

_п

= 1.5 㨪 4.0

Blow down type

Test section 䋺 600 mm 㬍600 mm

Transonic Supersonic

Internal force balance

Sting

Transonic and Supersonic Wind Tunnel

㪪㫀㪾㫅㫀㪽㫀㪺㪸㫅㫋㩷㪻㫀㪽㪽㪼㫉㪼㫅㪺㪼㫊㩷㫆㫅㩷㪸㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪺㪿㪸㫉㪸㪺㫋㪼㫉㫀㫊㫋㫀㪺㫊㪃㪿㫀㪾㪿㪼㫊㫋㩷㪣㫀㪽㫋㩷㪽㫆㫉㩷㫋㫉㫀㪸㫅㪾㫃㪼㩷㫄㫆㪻㪼㫃㪅

㪪㫈㫌㪸㫉㪼

㪚㫀㫉㪺㫃㪼

㪫㫉㫀㪸㫅㪾㫃㪼㪚

_㪣

㫍㫊㪘㫆㪘㩿㪤㪔㪇㪅㪊㪀

㪚

_㪛

㫍㫊㪘㫆㪘㩿㪤㪔㪇㪅㪊㪀

㪣㪆㪛㫍㫊㪘㫆㪘㩿㪤㪔㪇㪅㪊㪀

Wind tunnel test result (subsonic)

(10)

㪚

_㪣

㫍㫊㪘㫆㪘㩿㪤㪔㪇㪅㪐㪀

㪚

_㪣

㫍㫊㪘㫆㪘㩿㪤㪔㪈㪅㪊㪀

㪚

_㪛

㫍㫊㪘㫆㪘㩿㪤㪔㪇㪅㪐㪀

㪚

_㪛

㫍㫊㪘㫆㪘㩿㪤㪔㪈㪅㪊㪀

㪣㪆㪛㫍㫊㪘㫆㪘㩿㪤㪔㪇㪅㪐㪀

㪣㪆㪛㫍㫊㪘㫆㪘㩿㪤㪔㪈㪅㪊㪀

Wind tunnel test result (transonic)

㪚 _㪣㪄㱍㪺㫌㫉㫍㪼㫊㪃

㪤

_㺙

㪔㩷㪋㪅㪇

㪚 _㪛㪄㱍㪺㫌㫉㫍㪼㫊

㪚㪸㫅㫅㫆㫋㩷㫆㪹㫊㪼㫉㫍㪼㩷㫋㪿㪼㩷㫊㫀㪾㫅㫀㪽㫀㪺㪸㫅㫋㩷㪻㫀㪽㪽㪼㫉㪼㫅㪺㪼㫊㩷㫆㫅㩷㪸㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪺㪿㪸㫉㪸㪺㫋㪼㫉㫀㫊㫋㫀㪺㫊㩷㪺㫆㫄㫇㪸㫉㪼㪻㩷㫋㫆㩷㫊㫌㪹㫊㫆㫅㫀㪺㩷㫉㪼㪾㫀㫆㫅

Wind tunnel test result (supersonic)

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㪦㫀㫃㩷㪝㫃㫆㫎㩷㪭㫀㫊㫌㪸㫃㫀㫑㪸㫋㫀㫆㫅㩷㩿㪤㪔㪋㪅㪇㪃㩷㪘㫆㪘㪔㪊㪇㪻㪼㪾㪀

㪦㫀㫃㩷㪝㫃㫆㫎㩷㪭㫀㫊㫌㪸㫃㫀㫑㪸㫋㫀㫆㫅㩷㩿㪤㪔㪇㪅㪊㪃㩷㪘㫆㪘㪔㪊㪇㪻㪼㪾㪀

㪪㪿㫆㫌㫃㪻㩷㪹㪼㩷㪺㫃㪸㫉㫀㪽㫐㩷㫋㪿㪼㩷㪼㪽㪽㪼㪺㫋㩷㫆㪽㩷㫋㪿㪼㩷㫊㪼㫇㪸㫉㪸㫋㪼㪻㩷㫍㫆㫉㫋㫀㪺㪼㫊㫆㫅㩷㫋㪿㪼㩷㪸㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪽㫆㫉㪺㪼㫊

㪲㪫㫆㫇㩷㪭㫀㪼㫎㪴

Wind tunnel test result (flow pattern)

㪞㫆㫍㪼㫉㫅㫀㫅㪾㩷㪜㫈㫌㪸㫋㫀㫆㫅㪊㪛㩷㪝㫌㫃㫃㩷㪥㪸㫍㫀㪼㫉㪄㪪㫋㫆㫂㪼㫊㩷㪜㫈㫌㪸㫋㫀㫆㫅㩷㩿㪩㪘㪥㪪㪀㪚㫆㫅㫍㪼㪺㫋㫀㫍㪼㩷㫋㪼㫉㫄㫊㪘㪬㪪㪤㪄㪛㪭㫊㪺㪿㪼㫄㪼

㪭㫀㫊㪺㫆㫌㫊㩷㫋㪼㫉㫄㫊㪉㫅㪻㩷㫆㫉㪻㪼㫉㩷㪺㪼㫅㫋㫉㪸㫃㩷㪻㫀㪽㪽㪼㫉㪼㫅㪺㪼㪫㫀㫄㪼㩷㫀㫅㫋㪼㪾㫉㪸㫋㫀㫆㫅㪜㫌㫃㪼㫉㩷㪜㫏㫇㫃㫀㪺㫀㫋㩷㫄㪼㫋㪿㫆㪻

㪫㫌㫉㪹㫌㫃㪼㫅㪺㪼㩷㫄㫆㪻㪼㫃㫂㪄㱥㫋㫎㫆㩷㪼㫈㫌㪸㫋㫀㫆㫅㩷㫄㫆㪻㪼㫃

Numerical analysis : scheme

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㪝㫉㪼㪼㩷㪪㫋㫉㪼㪸㫄㩷㪤㪸㪺㪿㩷㪥㫌㫄㪹㪼㫉㪇㪅㪉㪐

㪫㫆㫋㪸㫃㩷㪧㫉㪼㫊㫊㫌㫉㪼㪈㪅㪌㪈㩷㬍㪈㪇 ^㪌㩷㪧㪸

㪫㫆㫋㪸㫃㩷㪫㪼㫄㫇㪼㫉㪸㫋㫌㫉㪼㪉㪐㪊㪅㪉㩷㪢

㪘㫅㪾㫃㪼㩷㫆㪽㩷㪘㫋㫋㪸㪺㫂㫊㪊㪇㩷㪻㪼㪾

㪦㫍㪼㫉㫍㫀㪼㫎

㪞㫉㫀㪻㩷㪧㫆㫀㫅㫋㫊㩷㪑㩷㪈㪍㪎㩷㬍㪎㪇㩷㬍㪌㪐

Flow condition and Computational Grid

75% 100%

25% 50%

㪘㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪝㫆㫉㪺㪼㫊㪜㫏㫇㪅㩷㪑㩷㩷㪚

_㪣

㪔㪇㪅㪈㪋㪃㩷㪚

_㪛

㪔㪇㪅㪈㪇㪚㪝㪛㩷㪑㩷㩷㪚

_㪣

㪔㪇㪅㪈㪋㪃㩷㪚

_㪛

㪔㪇㪅㪈㪉

㩿㪚㫀㫉㪺㫃㪼㩷㪝㫌㫊㪼㫃㪸㪾㪼㪃㩷㪘㫆㪘㪔㪊㪇㪻㪼㪾㪀

㪞㫉㫆㫎㫋㪿㩷㫆㪽㩷㫍㫆㫉㫋㫀㪺㪼㫊㩷㩽㩷㫊㫌㫉㪽㪸㪺㪼㩷㫇㫉㪼㫊㫊㫌㫉㪼

Comparison of CFD and Flow visualization

䊶㪪㫄㫆㫂㪼㩷㪑㩷㪊㪅㪌㩷㫄㪆㫊

㩿㩷㪩㪼㪔㪏㪅㪍㬍㪈㪇^㪋㩷㪀㪃

㪚㪝㪛㩷㪑㩷㪤㩷㪔㩷㪇㪅㪊㩷

㩿㩷㪩㪼㪔㪊㪅㪉㬍㪈㪇^㪍㩷㪀

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75% 100%

50%

㪘㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪝㫆㫉㪺㪼㫊 25%

㪜㫏㫇㪅㩷㪑㩷㩷㪚

_㪣

㪔㪇㪅㪈㪐㪃㩷㪚

_㪛

㪔㪇㪅㪈㪋㪚㪝㪛㩷㪑㩷㩷㪚

_㪣

㪔㪇㪅㪉㪉㪃㩷㪚

_㪛

㪔㪇㪅㪈㪎

㩿㪪㫈㫌㪸㫉㪼㩷㪝㫌㫊㪼㫃㪸㪾㪼㪃㩷㪘㫆㪘㪔㪊㪇㪻㪼㪾㪀

㪞㫉㫆㫎㫋㪿㩷㫆㪽㩷㫍㫆㫉㫋㫀㪺㪼㫊㩷㩽㩷㫊㫌㫉㪽㪸㪺㪼㩷㫇㫉㪼㫊㫊㫌㫉㪼

Comparison of CFD and Flow visualization

㪌㪇㩼㪉㪌㩼㪈㪇㪇㩼㪎㪌㩼

75%

100%

25% 50%

㪘㪼㫉㫆㪻㫐㫅㪸㫄㫀㪺㩷㪝㫆㫉㪺㪼㫊㪜㫏㫇㪅㩷㪑㩷㩷㪚

_㪣

㪔㪇㪅㪊㪇㪃㩷㪚

_㪛

㪔㪇㪅㪈㪍㪚㪝㪛㩷㪑㩷㩷㪚

_㪣

㪔㪇㪅㪊㪈㪃㩷㪚

_㪛

㪔㪇㪅㪈㪏

㩿㪫㫉㫀㪸㫅㪾㫃㪼㩷㪝㫌㫊㪼㫃㪸㪾㪼㪃㩷㪘㫆㪘㪔㪊㪇㪻㪼㪾㪀

㪞㫉㫆㫎㫋㪿㩷㫆㪽㩷㫍㫆㫉㫋㫀㪺㪼㫊㩷㩽㩷㫊㫌㫉㪽㪸㪺㪼㩷㫇㫉㪼㫊㫊㫌㫉㪼

Comparison of CFD and Flow visualization

㪉㪌㩼㪌㪇㩼㪎㪌㩼

㪈㪇㪇㩼

(14)

Space Transportation Systems Engineering Laboratory, Kyushu University CFD Experiment

Stream lines on Upper Surface

㪣㪸㫉㪾㪼㫊㫋㩷㫃㫆㫎㩷㫇㫉㪼㫊㫊㫌㫉㪼㩷㫉㪼㪾㫀㫆㫅

Circle Square Triangle Circle Square Triangle Stream lines on Upper Surface Pressure Contour on Upper Surface

Requirement for CFD

䊶higher accuracy for estimation of flow separation, aerodynamic forces, etc.

䊶reduction of computational time

Requirement for EFD

䊶improvement of wall interference correction, base drag correction, etc.

䊶spatial measurement

䊶non-contact measurement

(15)

Background .. again

Courtesy of JAXA

• In designing Reusable Launch Vehicle (RLV)

㸢 Aerodynamic Heating at reentry and supersonic flight

• Important Problems

– Aerodynamic Heating at stagnation point – Increase of Aerodynamic Heating

by transition of boundary layer

Necessity of Aerodynamic Heating Reduction

Example of TPS

Liquid

Gas

Transpiration cooling Film cooling

Heat absorption

Heat-resistant material

Passive method

Mass transfer cooling

Active method

Ablation

Intermediate method

Mechanical method

Cooling by opposing jet

Nose region

Courtesy of JAXA

(16)

The governing parameters

– The Mach numbers of free stream and jet – The diameter of the jet orifice

– The ratio of the total pressure of jet to that of free stream, PR

f 0

0

p

PR p

^j the ratio of total pressure of opposing jet

to that of free stream

The Flow Field of the Opposing Jet

z

Measurement of the heat flux

z

Visualization of the flow field with Schlieren method

z

Kyushu univ. supersonic wind tunnel

Experimental Research

• Free stream (average on exp.䋩

Mach Number 3.96

Stagnation Pressure 1.37 MPa Stagnation Temperature 397 K, 497 K

Reynolds Number 2.1x10

⁶

• Secondary flow

Mach Number 1.0

Total Pressure Ratio, PR 0.0 - 0.8

Stagnation Temperature 300 K

(17)

Experimental model : blunt body

• The calorimeter gauges are attached at 20 to 90 degrees (every 10 degrees)

• The model is installed into the free stream after the flow becomes steady flow.

The diameter of the model : 50 mm

The diameter of the jet orifice : 4 mm

Mach number at the jet orifice : 䋱.0

Numerical Analysis Method

• Governing Equation: Reynolds averaged axisymmetric Navier-Stokes equation (RANS)

• Time Integration : LU-ADI method

• Convection Term: AUSM-DV scheme with MUSCL

interpolation

• Viscous Term: 2nd-order central difference scheme

• Turbulence model : k-㱥 two equation model

– C_㱘term is introduced in order to prevent the excessive generation of kin collision region. Craft et.al(1996).

(18)

Flow Conditions and Grid for supersonic flow

Mach number 3.96

Total pressure [MPa] 1.37 Total temperature [K] 397

Mach number 1.0

Pressure ratio 0, 0.4, 0.6, 0.8 Total temperature [K] 300

Wall temperature [K] 295

• Grid (242 㬍 160)

Diameter of body [mm] 50 Diameter of jet orifice [mm] 4

Flow Conditions and Grid for hypersonic flow

Mach number 8.0

Total pressure [MPa] 4.5 Total temperature [K] 800

Mach number 1.5

Pressure ratio 0.0251~0.0859 Total temperature [K] 300

Wall temperature [K] 300

• Grid (240 㬍 160)

This flow conditions and grid configuration are based on the experiment conducted by Tokyo Univ. in 1975.

Diameter of body [mm] 40 Diameter of jet orifice [mm] 4.34

(19)

Flow field at each PR (Mach 3.96)

No jet PR = 0.40 PR=0.60 PR=0.80 Calculated

Experiment

No jet PR=0.40 PR=0.60 PR=0.80 EXP

CFD

No jet PR=0.40 PR=0.60 PR=0.80

q (W/m )w2

^(degrees)

0 20 40 60 80

-20000 0 20000 40000 60000 80000 100000 120000 140000 160000 180000

Heat flux distribution of CFD is very similar to that of the experiments

Heat flux at each angles decreases as PR

increases

Boundary layer transition occurs

Heat flux decreases in recirculation region Heat flux increases

due to recompressed shock

Comparison of Heat Flux (Mach 3.96)

(20)

Flow field change by PR (Mach 3.96)

No jet PR=0.40 PR=0.60 PR=0.80 EXP

CFD No jet PR=0.40 PR=0.60 PR=0.80

q (W/m )w2

(degrees)

0 20 40 60 80

-20000 0 20000 40000 60000 80000 100000 120000 140000 160000 180000

Temperature contour

PR = 0.40 PR = 0.80

Heat flux decreases in recirculation region Heat flux increases due to recompressed shock

PR=0.0251 PR=0.0549 PR=0.0859

Calculated

Experiment

Flow Field at each PR (Mach 8)

(21)

Pressure Distributions (Mach 8)

Pressure Distributions

CFD pressure distribution shows good agreement with experimental measurement.

Heat Flux Distributions (CFD)

Heat flux decreased more considerably than the case for supersonic.

Heat flux at each angles decreased as PR increases.

• High density in stagnation region

• Low temperature jet

Large heat capacity

• Low temperature recirculation region

Mach number Density

Temperature

• Remarkable reduction of aerodynamic heating Mach 8

PR=0.0859

Understanding the flow field

(22)

Consideration

• The opposing jet is useful to reduce aerodynamic heating in supersonic and hypersonic flow.

• To understand the mechanism of reducing aerodynamic heating by the opposing jet, detailed flow field should be clarified.

• CFD is very powerful tool to understand the flow field, but has to be validated.

Background .. Once again

Development of scram-jet engine is now in progress as a

propulsion system of hypersonic transports and space planes.

䊶In scram-jet engine, the speed of air is very high.

Hence rapid mixing and combustion of air and fuel is required. (Air residence time in combustor is 10

^-3

䌾10

^-4

sec)

䊶At high Mach number, suppression of development of

shear layer occurs and it makes mixing of air and fuel

difficult .

(23)

Objective of the present study

Investigation of the effect of the injection angle Ǫ of three-dimensional circular

nozzle on supersonic mixing flow

Free Stream

Injection Angle(㱎)

Schematic of experimental facility

Air reservoir

Compressor 65m

³

30kgf/cm

²

Helium cylinder

Traverse equipment Test section

Flat plate modelSampling probe

Sampling chambers

(24)

Schematic of the experimental model

y

x

150

130

300 z x

520

24 Mount

㱢3

Free stream

Test section

7.5㫦

Secondary gas

Experimental conditions

Free stream

Secondary gas

Gas

Mach number

Total temperature Gas

Mach number Total pressure Total temperature

Air 3.76 1. 12 MPa

286.9 K

286.9 K Helium

1.0 0.40 MPa

Total pressure

(25)

Flow visualization by Schlieren method

(a)㱎䋽30㫦

(b)

㱎䋽90㫦

(c)

㱎䋽150㫦

䊶separation shock wave 䊶bow shock wave

As injection angle 㱎 becomes large, separation region becomes wider and bow shock wave becomes stronger.

Numerical method

Governing equations : Reynolds averaged 3D full N-S Convective terms : AUSM-Plus scheme

Viscous terms : 2nd order central difference Time integration : LU-ADI method

Turbulence model : k-㱥 two equation model

with Low Reynolds number effect

number of grid point : 116㬍42㬍75

(26)

Stream line on surface

(a)

㱎=30㫦

(b)

㱎= 90㫦

(c)

㱎= 150㫦

Free Stream

Injection Angle (㱎)

Wall pressure distribution at y=0

(a)

㱎=30㫦

(b)

㱎= 90㫦

Calculation Experiment

(c)

㱎= 150㫦

(27)

Volume fraction distribution

Calculations

㱎=30deg

㱎=90deg

㱎=150deg

Experiment

X=0.04m

Mixing efficiency and Total pressure loss

Mixing efficiency

³

A H

dA uf ĭ dA uf x

2 2

U U K

¯®

! d

) 1 ) , , ( if ( ) , , (

) 1 ) , , ( if ( 1

z y x z

y x

z y ĭ x

I I

I

2 2

35 1 ) , , (

H H

f z f

y

x u

I

Stoichiometric mass ratio ... Hydrogen : Air = 1 : 35 A : Cross section of test section

: Mass fraction of hydrogen

: Local equivalent ratio in the cross section dA

H2

If

dA up dA

up

dA up x

j

i A t

A t

x

A t

³

U U

S( ) 1 ⁽ ⁾U

p_t: Local total pressure A : Cross sectional area at x

A_i: Cross sectional area at the entrance of the test section A_j: Cross sectional area of the injector

(28)

EFD/CFD Activities in Research for Reusable Launch Vehicles