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American Institute of Aeronautics and Astronautics

AIAA 2002-4192

CFD Validation with MeasuredTemperatures and Forces for Thrust

Vector ControlP-A. Rainville, A. deChamplainand D. KretschmerUniversit Laval, Qubec

Qubec, G1K-7P4 Canada

R. Farinaccio and R.A. Stowe

Defence R&D Canada - Valcartier, Qubec

Qubec, G3J-1X5 Canada

38th AIAA/ASME/SAE/ASEEJoint Propulsion Conference & Exhibit

7-10 July 2002Indianapolis, Indiana

38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit7-10 July 2002, Indianapolis, Indiana

AIAA 2002-419

Copyright 2002 by the author(s). Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

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AIAA-2002-4192

1American Institute of Aeronautics and Astronautics

CFD VALIDATION WITH MEASURED TEMPERATURES AND FORCES

FOR THRUST VECTOR CONTROL

P-A. Rainville*, A. deChamplain, D. Kretschmer

Universit Laval

Qubec, QCG1K-7P4 Canada

R. Farinaccio, R.A. Stowe

Defence R&D Canada ValcartierQubec, QC

G3J-1X5 Canada

Abstract

The objective of this work was tocalculate and validate a transient simulation

of a TVC nozzle system. The calculationswere done on a structured mesh with the

commercial code Fluent. The initial steadystate calculations gave valuable results and

showed that compressible and viscous effects,along with internal heat conduction, were

modelled correctly. Particular attention wasalso paid to the convergence rate. For the

unsteady calculations, the experimentalvalidation of the temperature is quite

reasonable near the steel insert holding thevane to the nozzle assembly, but much lower

in the CIT portion of the vane. The validationof the experimental thrust-time curve is also

good initially but deteriorates somewhattowards the end of motor firing after two

seconds of operation. These preliminaryresults are encouraging, but more refinements

are still necessary to account for the twodifferent types of material and their properties

inside the vane.

Introduction

Over the last twelve years, onlylimited work has been done to determine the

temperature distribution in the vanes of athrust vector control (TVC) system.

Danielson1

completed an experimental studyto estimate the temperature distribution and

the rate of erosion on jet vanes used for TVC,but he did not actually measure temperature

distribution in the vanes. Rahaim et al.2

calculated the flowfield around the vane with

a 2D fluid simulation code, but they did notsimulate the flow in the entire nozzle.

However, the thermal conduction within thevane was calculated, albeit with a different

code. Building on previous attempts with asteady state simulation

3, the present work is

to calculate the unsteady flow field in thenozzle in three dimensions and to include

thermal, compressible, and viscous effects.

The objective of this study is tovalidate the commercial code Fluent for the

simulation of the unsteady flow field withinthe nozzle of a solid propellant motor

equipped with TVC. The purpose of a TVCsystem is to allow directional control of a

flight vehicle with jet vanes inside the motornozzle acting on the rocket exhaust plume.

The experimental data for the validation ofFluent are based on time-dependent test

results that were completed at Defence R&DCanada Valcartier (DRDC Valcartier).

__________________________________________________________________

* Graduate Student, Mechanical Engineering

Professor, Mechanical Engineering, Member AIAA

Scientist, Propulsion Group, Member AIAA Scientist, Propulsion Group, Senior Member AIAA

Copyright 2002 by the Department of National

Defence and Universit Laval, Canada.Published by the American Institute of Aeronautics andAstronautics, Inc., with permission.

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These experimental results include several

parameters for the solid propellant motor thatestablish the operating conditions for the

numerical simulation of the TVC system.

The experimental data for thevalidation include the time-dependent forces

measured on the TVC vanes and the time-dependent temperatures measured inside the

vanes during the motor firings. Because themotor has a progressive thrust-time profile,

unsteady implicit calculations were done withthe motor pressure varying with time over the

4-second burn.

Geometry

To reduce the geometry to a morereasonably sized grid, the calculations were

done only for the nozzle section of the rocketmotor. The drawing of this nozzle is shown

on Figure 1. The combustion chamber of themotor is situated to the left of the nozzle.

Since the combustion chamber is notmodelled, the flow into the nozzle is

simulated with a pressure inlet boundarycondition, with this pressure being the

pressure measured in the combustion chamberof the rocket motor. Figure 2 shows the

pressure-time profile of the motor.

The two pairs of vanes providingthrust vector control were placed near the exit

plane of the nozzle to control roll, pitch, andyaw by deflecting the rocket exhaust plume.

The vanes were made of steel or copper-infiltrated tungsten (CIT). The temperatures

were measured with three thermocouples. Asillustrated in Figure 3, the first one (point 1)

was placed in the vane at 2,54 cm (1 in) fromthe base edge of the vane. The second

(point 2) was also inside the vane, but at1,25 cm (0,5 in) from the base edge. The last

one (point 3) was fixed in the pedestal of thevane at minus 1,25 cm (0,5 in) from the vane

base edge. All three thermocouples werealigned with the axis of the vane and its

pedestal. The forces and moments in three

directions produced by the motor weremeasured with strain gages mounted on an

aerodynamic force balance.

Numerical Modelling

Based on previous work3, a structured

mesh was found to offer better results for

these 3D nozzle simulations. The mesh sizewas limited to approximately 675 000 cells

due to the capacity of the two-gigabyte PCcomputer memory and to give reasonable

calculation time. To improve the accuracy ofthe numerical solution near the vane, only a

90-degree sector of the geometry wasmeshed. The vane was placed at the centre of

the sector. Figure 4 shows a better view of themesh on and around the vane for the quarter

sector. The boundary is limited with arotational periodic condition. This boundary

condition was preferred to a symmetryboundary condition to account for transversal

flow that could exist in the nozzle.

Figure 5 shows the positions of theboundary conditions. At the nozzle entrance,

a pressure inlet (blue) is applied. The totalpressure is fixed with the pressure as

measured in the combustion chamber of themotor. The Chemical Equilibrium

Applications (CEA) software from NASA4

was used to calculate the imposed

corresponding total temperature. Theseprofiles are then approximated with a time-

wise polynomial function. Figure 2 shows thetotal pressure profile and Figure 6 the total

temperature profile.

Around the nozzle, a pressure far fieldboundary condition (pink) is applied. It

simulates the Mach 3 nozzle motion atatmospheric pressure. The static temperature

is 300 K and the total temperature is 840 K. Apressure outlet (red) closes the calculation

domain. It was chosen to reduce shock wavereflections. The far field domain behind and

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around the nozzle is just big enough to

eliminate any boundary effect around thevanes.

The mesh was generated using the

commercial codes Gambit

and Gridgen

,while Fluent was used to perform the flow

field simulation with a Reynolds averagedcompressible Navier-Stokes (RANS) solver.

The RANS equations were closed with the

k- RNG turbulence model, and the flow near

the solid walls was predicted with a standardwall function. The working fluid is air with

compressible properties evaluated using theperfect gas law. The wall boundary condition

is coupled for thermal interaction between the

fluid and the solid surfaces. Calculations weredone with a transient implicit solver.

With an effort to reduce thecalculation time to reasonable values, a

special strategy was developed. The first stepwas to complete a steady state calculation.

After about a week a decision can be made asto whether or not the flow is correctly

simulated. If it is, then the steady solution isused as a starting point for the transient

calculation, and overall convergence can beachieved more quickly. The solid wall

temperature was set to the original value of300 K. The calculation, over a time interval

of four seconds, is then run with a 0,05second time step. This part of the work would

take up to few weeks to complete.

Results

The steady state solution appears to

give a reasonable estimate of the flow field.Figure 7 shows the pressure contour in a

plane cutting through the middle of the vane.The diamond-shaped shock wave pattern is

easy to see and seems to be consistent withgas dynamic theory. Figure 8 shows the Mach

number contour in the same plane as thepressure contour. It is interesting to see a

fairly low Mach number at the rear of the

vane since the plane cuts through the low

speed wake behind the vane. Figure 9 showsanother view of the pressure contour. This

figure is taken in a plane perpendicular to theplane shown on Figure 7 or 8, at

approximately the middle point of the vane.This view helps to see the complex shock

wave pattern that results from the strong gasdynamic interaction between the nozzle wall

and the vane. The black line around thecontour gives an idea of the nozzle position.

But even if flow structure is adequate, thevalidation is not satisfactory with the steady

state results, especially with the temperaturefield inside the vane reaching the same

temperature as the hot gas at 3 100 K asshown on Figure 10. In actual fact, even after

the 4-second motor firing, it will be shown inthe next paragraph the vane temperature does

not exceed much the 2 000 K. A comparisonwith the transient experimental results will

therefore provide a more suitable assessmentof the time dependence.

With a plane cutting through the

middle of the vane, Figure 11 presents theinternal temperature field for the unsteady or

time-dependent calculation. Since the internalheat conduction is a relatively slow process,

even after two seconds of motor operation theinside of the vane is still much cooler than the

surrounding hot gases which have astagnation temperature of 3 100 K. The

temperature validation is done only with theCIT vane because the vane made out of steel

lost too much of its shape due to erosion5, and

CFD calculations could not account for the

erosion phenomena and change in shape. TheCIT vane shape did not change significantly,

which should not affect too much the finalresults. Figures 12 and 13 present the

temperature inside the vane as a function oftime. The CIT vane material is defined with

cold constant properties in Figure 12. A goodagreement with experimental data is possible

only for the temperature measurement in thevane pedestal. These results would suggest

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that something peculiar happens in the CIT

material during the motor firing. From ananalysis of the experimental data, it is

suspected that the copper infiltrated into thetungsten would evaporate at high temperature

with an appreciable change in the materialproperties. While more effort must be done to

model this properly and achieve better results,it has not been possible to date. As an

example, Figure 13 presents results forcalculation done with the steel vane at

constant properties. Curiously, the steelproperties offer a better agreement with the

experimental data than for the CIT vane. Thesteel insert in the vane can explain this good

agreement for the second and third points.The second measurement position is very

close to the end of the steel insert. During allcalculations, the third measurement point, in

the steel pedestal, is calculated with thematerial properties of the steel. For the first

point closer to the tip of the vane, it wouldappear that a solution in between these two

simulations would be satisfactory. To achievethis, it would appear that the grid for the vane

has to be redone to include separate blocks tocharacterize separately the different materials

in the vane. Consequently the insert and thepedestal could be simulated with steel

properties and the vane with CIT properties.Possibly, these properties will have to be

temperature dependent as well.

The second part of the validation wasdone with force measurements. Figure 14

presents the thrust as a function of time forthe motor. The results are quite good for the

first two seconds. However, there are someinconsistencies with the experimental data

that need further explanations at this point.The simulation work is done with a four-vane

system at a 0 angle-of-attack. The

experimental data for the 8084 motor with

four vanes had two of them at a 10 angle-of-

attack. Unfortunately all four steel vanes were

sheared off during the run explaining the

increase in thrust at around 2,7 s in the curve

for the 8084 motor on Figure 14. The 8100

motor only had two vanes with no angle-of-

attack. The thrust curve for the 8100 motor is

higher because the burning pressure is

slightly higher. This difference between the

experimental data and the simulation couldalso be partly explained with more erosion at

the nozzle throat for which the simulation

could not account. Another factor to consider

is that the fluid properties used were for air

whereas they should be changed to reflect

more closely exhaust gas properties as they

exist in the nozzle. Considering these factors

in future simulations should provide more

accurate convergence and should improve the

agreement between numerical and

experimental results.

Conclusions

In the present article, steady and

transient calculations were done on a TVC

nozzle system. The boundary conditions were

based on the experimental data for a more

realistic simulation. Calculations were done

in two parts, a steady calculation to obtain the

main flow structure, and a transient part to

validate the time dependent force and

temperature profiles. The steady state solution

gave a fairly good flow structure. The

transient temperature validation was

reasonable near the steel insert in the vane.

However, the heating process in the CIT

should be modeled with a remeshing of the

vane to allow different thermal properties for

the two materials inside the vane, and also

allow these properties to vary with

temperature. The agreement with the

experimental thrust curve is quite good at the

start but became low during the latter part. A

better validation should likely result with the

use of properties for the combustion products

rather than using the properties for air. With

new improvements in the simulation code

Fluent, the erosion of the vanes could also be

modeled in future work.

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Acknowledgements

Funding for this study was provided by

Honeywell (formerly Allied Signal), DefenceR&D Canada (DRDC), the Natural Sciences

and Engineering Research Council of Canada(NSERC), Fond de Recherche sur la Nature et

les Technologies Qubec (FCAR) andUniversit Laval.

References

1. Danielson, A., Inverse Heat TransferStudies and the Effects of PropellantAluminium on TVC Jet Vane Heating andErosion, AIAA-90-1860.

2. Rahaim, C.P., Cavalleri, R.J.,McCarthy, J.G., Kassab, A.J., Jet VaneThrust Vector Control: a Design Effort,AIAA-96-2904.

3. Rainville, P.A., deChamplain, A.,Kretschmer, D., Hamel, N., Farinaccio, R.,Stowe, R., Simulation Numrique d'uneTuyre Mach 3 avec Ailettes pour leContrle du Vecteur Pousse, Ve Colloque

Interuniversitaire Franco-Qubcois,Thermique des systmes, diteur J. Brau,Lyon, Mai 2001.

4. McBride J., Gordon S., CompleteProgram for Calculation of ComplexChemical Equilibrium Compositions andApplications, NASA Reference Publication1311, June 1996.

5. Harrisson V., de Champlain A.,Kretschmer D., Farinaccio R., Stowe R.,Optical Technique To Quantify Erosion OnJet Vanes For Thrust Vector Control, AIAA-2002-4090.

Figure 1 Nozzle geometry with two of the four vanes assembly near the exit plane.

Combustion

chamber

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Figure 2 Combustion chamber pressure-time profile imposed as the inlet pressure of thenozzle for a time-dependent numerical solution.

Figure 3 Experimental points for temperature measurements inside the vane

Figure 4 Meshed vane as mounted at the exit plane of the nozzle.

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Figure 5 Complete view of the mesh with the various boundary conditions: a pressure inlet

(blue), a pressure far field (pink), a pressure outlet (red), nozzle walls (black).

Figure 6 Combustion chamber temperature-time profile imposed with the inlet pressurecondition of the nozzle for a time-dependent numerical solution.

Figure 7 Pressure contours in Pa at the mid-plane of the vane for the steady state

calculation.

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Figure 8 Mach number contours at the mid-plane of the vane for the steady statecalculation.

Figure 9 Pressure contours in Pa for a perpendicular plane cutting through the vane for thesteady state calculation.

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Figure 10 Temperature contours in K at the mid-plane of the vane for the steady state

calculation.

Figure 11 Temperature contours in K at the mid-plane of the vane for the transientcalculation after two seconds of motor firing.

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Figure 12 Internal vane temperature-time profile comparing numerical simulation (solid line)

with experiment (dotted line) for the CIT vane at constant properties.

Figure 13 Internal vane temperature-time profile comparing numerical simulation (solid line)

with experiment (dotted line) for the steel vane at constant properties.

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Figure 14 Motor thrust from numerical simulation (solid line) and experiment (dotted line).

aiaa-2002-4192-430

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