Derivation of Preliminary Ascent Vibro-Acoustic Environments for the Crew Exploration Vehicle...

Derivation of Preliminary Ascent Vibro-Acoustic Environments for the Crew Exploration Vehicle

Presented by:Mike Yang / ATA Engineering

Nancy Tengler / Lockheed Martin Corporation

atSpacecraft and Launch Vehicle Dynamic Environments Workshop

June 27-29, 2006

2

Presentation Outline

• Description of CEV• Development of aeronoise model for different flow regimes

– Attached turbulent boundary layer– Compression corner– Expansion corner

• Comparison to existing empirical data– Titan IV flight data– Apollo/Saturn wind tunnel data

• Prediction of Launch Abort System (LAS) motor noise• Prediction of internal responses using SEA/FEA

– Comparison of SEA and FEA responses• Summary and Conclusions

3

CEV is Comprised of Several Modular Parts

Crew Module Service Module

Spacecraft Adapter

Launch Abort System

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Objectives

Part 1• Predict external acoustic environment for CEV during the liftoff,

nominal ascent, and abort events– Aeronoise– LAS Abort Motor

Part 2• Use derived external environments to predict internal responses

– Acoustic cavities– Panel responses

• Use predicted responses to aid in design evaluation and refinement

5

Cp Contour Plots Reveal Different Flow Regimes

Attached TBL

Compression Plateau Compression Peak

Expansion Peak Expansion Plateau Attached TBL

q

PPC staticp

tCoefficien Pressure Static

6

Aeronoise Models are Comprised of Three Components

RMS Cross-spectraAutospectra

Increasing Uncertainty

Derived loads are dependent on:

• Dynamic Pressure (q)• Mach (M)• Atmospheric Properties (altitude)

• Density ()• Speed of sound (c)• Ratio of specific heats ()• Kinematic viscosity ()

Flight Parameters Spacecraft Geometry• Determines flow regime• Affects RMS Pressure for

some flow regimes

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RMS Pressure is Typically Expressed as a Fluctuating Pressure Coefficient

• K, A = Constants determined from experimental data• F = A function of Mach number• P2/P1 = Pressure ratio across shock wave (Function of mach, ratio of

specific heats, shock wave angle)

q

PC RMSp

Cp Equations

Plateau Peak

Attached TBL

Compression Regime

Transonic

Supersonic

Expansion Regime

F

KTBL

2

,

1 M

K pltrncomp

F

K

P

P comp sup,

1

2

F

K

P

PA

P

PAA comp sup,

2

1

23

1

221

2

exp

1 M

K pl

exp42

exp

1A

M

K pk

2

,

1 M

K pktrncomp

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Fluctuating Pressure Coefficient as a Function of Mach(ISS Trajectory Used for Atmospheric Properties)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.50 1.00 1.50 2.00 2.50 3.00 3.50 4.00 4.50 5.00

Mach

Del

ta C

p

Turbulent Boundary Layer Compression Shock - Plateau Compression Shock - Peak

Expansion Shock - Plateau Expansion Shock - Peak

Peak Expansion and Compression Shock Regimes have Highest Cp Values

Peak Expansion Shock

Peak Compression Shock

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Noise Spectrum Due to Compression Corner

135.0

140.0

145.0

150.0

155.0

1 10 100 1000 10000

Frequency (Hz)

SP

L (

dB

)

Peak

Plateau

Shape of Autospectra Curve is Function of a Parameter which is Regime-Dependent

• C is regime-dependent, and shifts curve to the left

• Generally:

– Cpeak > Cplateau

– Ccomp > Cexp > CTBL

• Slope of 1/3-octave band spectrum:– -10 dB/dec. at low freqs.– +10 dB/dec. at high freqs.

• Function always integrates to 1

+10 dB/decade -10 dB/decade

Increasing C

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Default VA-One Cross-Correlation Coefficients were Used

• Default values are c()=.1, c()=.72, ()=1, ()=0, =0

kekeR ii

kkckkc

coscos,,

2

*22

2

*22

3

1

3

1

'CU

k

'CU

k UUC '

Along Flow Cross Flow

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

0 0.5 1 1.5 2

Distance

Cro

ss-C

orr

elat

ion

Co

effi

cien

t

Along Flow

Across Flow

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Different Geometries Used to Verify Aeronoise Model

Coe, C.F. Nute, J.B. "Steady and Fluctuating Pressures at Transonic Speeds on Hammerhead Launch Vehicles," NASA TM X-778, December 1962.

Jones, George W. Jr., and Foughner, Jerome T. Jr. "Investigation of Buffet Pressures on Models of Large Manned Launch Vehicle Configurations." NASA Technical Note D-1633. May 1963. p. 32.

Shelton, J.D. "Collation of Fluctuating Buffet Pressures for the Mercury/Atlas and Apollo/Saturn configurations." NASA CR 66059. p. 15.

Coe, Charlie F., and Kaskey, Arthur J. "The Effects of Nose Bluntness on the Pressure Fluctuations Measured on 15 degree and 20 degree Cone-Cylinders at Transonic Speeds." NASA TM X-779. January 1963. p. 7.

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Predicted Autospectra Anchored to Flight Data

125

130

135

140

145

150

155

10 100 1000 10000

Frequency (Hz)

120

125

130

135

140

145

150

155

160

10 100 1000 10000

Frequency (Hz)

120

125

130

135

140

145

150

10 100 1000 10000

Frequency (Hz)

Expansion Peak Expansion Plateau

Attached Turbulent Boundary Layer

SP

L (d

B)

SP

L (d

B)

SP

L (d

B)

• Derived aeronoise model envelopes majority of data

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Aeronoise RMS Levels Anchored to Wind Tunnel Data

Compression Regime with Modifications Made to ModelDC offset = 0.025

0.000

0.020

0.040

0.060

0.080

0.100

0.120

0.140

0.160

0.180

0.5 1 1.5 2 2.5Mach

De

lta C

p

Compression Shock - Plateau Compression Shock - Peak

CR 66059 - T1 CR 66059 - T2

8% Tower 3 Freon - Mic 1 8% Tower 1 air - Mic 1

8% Tower 1 Freon - Mic 1 8% Tower 2 freon - Mic 1

1.6% Tower 1 - Mic 1 1.6% Tower 2 - Mic 1

Apollo/Saturn-like Wind Tunnel Model

T1

T2

Shelton, J.D. "Collation of Fluctuating Buffet Pressures for the Mercury/Atlas and Apollo/Saturn configurations." NASA CR 66059. p. 15.

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NASA-SP-8072 Used to Predict LAS Abort Engine Noise

• “Method 2” divides the plume into slices– Each slice has a different sound power spectrum– SPL at CEV surface calculated by acoustically radiating sound back to

surface– An additional 1-3 dB were added to account for surface reflections

• This method assumes that there is no plume impingement.– How would we handle this if there was impingement?

Zone 9-4

Zone 9-3

Zone 9-2

Zone 9-1

Zone 8-2

Zone 8-1

Zone 7-2

Zone 7-1

Zone Center

LAS Motor Noise Spectrum

100

110

120

130

140

150

160

1 10 100 1000 10000

Frequency (Hz)

SP

L (r

ef =

20

mic

roP

a) Zone 7-1

Zone 7-2

Zone 8-1

Zone 8-2

Zone 9-1

Zone 9-2

Zone 9-3

Zone 9-4

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Internal Responses were Computed using SEA

• Models created in VA-One

• SEA is ideally suited for vibroacoustic predictions at high frequencies

• Symmetric half-model was used to reduce computation time.

Three models:

1. Liftoff

• DAF excitation

• Sea-level

2. Nominal Ascent

• TBL excitation

• Altitude ~ 15K feet

3. LAS Abort

• TBL and DAF excitation

• Altitude ~ 31K feet

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SEA Results Reveal Critical Events

• Crew Module– Highest noise levels occur during LAS Abort

• Spacecraft Adapter – Highest noise levels occur during Liftoff

Crew Module Cavity SPL

60.0

70.0

80.0

90.0

100.0

110.0

120.0

130.0

140.0

10 100 1000 10000

Frequency (Hz)

Ca

vit

y S

PL

, dB

(R

ef

= 2

0 m

Pa

)

Liftoff

Nominal Ascent

LAS Abort

Spacecraft Adapter Cavity SPL

60.0

70.0

80.0

90.0

100.0

110.0

120.0

130.0

140.0

10 100 1000 10000

Frequency (Hz)

Ca

vit

y S

PL

, dB

(R

ef

= 2

0 m

Pa

)

Liftoff

Nominal Ascent

SPL in Cavity 1 SPL in Cavity 2

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Hybrid FEM-SEA Model Used to Compute Low-Frequency Response of Thrust Cone

• VA-One’s hybrid analysis capability was used• “Sensors” were placed at several nodes on Cone FE subsystem

– Sensor response was averaged (spatial average)• Low-frequency response of Thrust Cone was verified

Thrust Cone Vibration During Nominal Ascent

1.0E-04

1.0E-03

1.0E-02

1.0E-01

1.0E+00

1.0E+01

10 100 1000

Frequency (Hz)

g^

2/H

z

SEA

Hybrid

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Summary & Conclusions

• Flow around a vehicle can be divided into flow regimes– These flow regimes will have different environments

• Aeronoise model consists of three parts:1. RMS Pressure

• Peak expansion and peak compression regimes are highest2. Autospectra

• Shape of spectrum is function of the flow regime3. Cross-spectra

• Decaying sine along flow, decaying exponential across flow• Aeronoise model was anchored to flight & experimental data

– Saturn/Apollo– Titan– Others (not shown)

• NASA-SP-8072 was used to predict LAS Abort Motor Noise– How can we predict noise due to plume impingement?

• SEA and Hybrid FEA-SEA analysis was used to predict internal CEV responses