n

ANALYSIS

OF ""Vf^

SSME HJEL PREBURWKR FL

Prepared For:

£NftSft-CE-178803) COMPUTATIONaL ANftLYSIS OF N86-23612

THE SSB£ FDEL PBEBDENEB FLOW Final Beport

. (Continuum, Inc.) 59 p HC flOa/HF &01

CSCL 21H Onclas

G3/20 16967

By:

Ter<-S?e Wi-ng

Hic'narc1' C. Farrier

COSTIN'JTOM. Inc.

4715 Univeio'ty Drive

Suitp 11B

Honir-viUc, M.

February 2fc, -9&1'
TABLE OF CONTENTS

Summary iv

Introduction 1

VAST Methodology

o The Transport Equations 4

o The Num erical Analog 5

o VAST Differencing Scheme 6

Turbulence Modeling 7

Results of the Single Coaxial Hydrogen/Air Jet Calculation 9

Combustion Chemistry 16

Subsonic Boundary Conditions

o Fixed Integrated Mass Flow Rate Boundary Conditions 20

o Injector Simulation Upstream Boundary Conditions 21

General Flow Features and Assumptions 23

Axisymmetric Modeling of the Fuel Preburner

o Grid Definition 27

o Boundary and Inlet Conditions 30

o Results for the Axisymmetric Fuel

Preburner Models 32

Three-Dim ensional Modeling of the Fuel Preburner

o Grid Definition 39

o Boundary and Inlet Conditions 42

o Results for the Three-Dim ensional Fuel

Preburner Model 44

Conclusions 52

Recommended Measurements Study 54

References 55
TABLE OP FIGURES

Fig. 1 Schematic of SSME Fuel Preburner 2

Fig. 2 Centerline Velocity Decay of a Coaxial H2/Air Jet 10

Fig. 3 Centerline Velocity Decay of a Coaxial H2/

Air Jet

From Ref. 8 11

Fig. 4 Comparison of Model Predictions of Lateral Hydrogen/

Air Mixing with Data at Different Axial Stations

(a) Velocity Profiles; x/d = 0.0 to 9.58 12

(b) Velocity Profiles; x/d = 15.4 to 63.6 13

(c) Hydrogen Mass Fraction Profiles;

x/d = 0.0 to 9.58 14

(d) Hydrogen Mass Fraction Profiles;

x/d = 15.4 to 63.6 15

Fig. 5 Temperature Histories for H2/Air at

One Atmosphere ( <t>= 1 ) 17

Fig. 6 Temperature Histories for H2/Air

 at

One Atmosphere ( 4>= 2 ) 18

Fig. 7 Temperature Histories for H2/O2 at <t>=7. 4 19

Fig. 8 The Layout of SSME Fuel Preburner Injector Faceplate 26

Fig. 9 Axisymmetric SSME Fuel Preburner Coarse Grid

(890 nodes) Definition 28

Fig. 10 Axisymmetric SSME Fuel Preburner Fine Grid

(5965 nodes) Definition 29

Fig. 11 Axisymmetric SSME Fuel Preburner Flow Vectors 33

Fig. 12 Axisymmetric SSME Fuel Preburner Temperature

(Coarse Grid) 34

Fig. 13 Axisymmetric SSME Fuel Preburner Hydrogen Mass

Fraction (Coarse Grid) 35

Fig. 14 Axisymmetric SSME Fuel Preburner Temperature

(Fine Grid) 36

Fig. 15 Axisymmetric SSME Fuel Preburner Inlet and

Exit Temperature 37

Fig. 16 Axisymmetric SSME Fuel Preburner Hydrogen Mass

Fraction (Fine Grid) 38

Fig. 17 Front View of the 3-D SSME Fuel Preburner Grid

Definition With 30-Deg. Angle 40

Fig. 18 Top View of the 3-D SSME Fuel Preburner Grid

Definition 41

-ii-
Fig. 19 The Swirling Flow Vector Component of the

Igniter Inlet 46

Fig. 20 3-D SSME Fuel Preburner Flow Vectors 47

Fig. 21 3-D SSME Fuel Preburner Temperature 48

Fig. 22 3-D SSME Fuel Preburner Cross-Sectional Temperature

at Bottom of Baffles 49

Fig. 23 3-D SSME Fuel Preburner Cross-Sectional Temperature

at 1.63 in. Above Turbopump 50

Fig. 24 SSME Fuel Turbine Inlet Temperature Profile at FPL 51

-in-
SUMMARY

The objective of this study was to develop a computational fluid dynamics model which

simulated the steady state operation of the SSME fuel preburner. Specifically, the model

will be used to quantify the flow factors which cause local hot spots in the fuel preburner

in order to recommend experiments whereby the control of undesirable flow features can

be demonstrated. This report presents the results of a two year effort to model the

preburner. In this effort, investigating the fuel preburner flowfield, the appropriate

transport equations were numerically solved for both an axisymmetric and a threedimensional configuration. Continuum's VAST (VAriational Solution of the Transport

equations) code, in conjunction with our CM-1000 Engineering Analysis Workstation and

the NASA/Ames CYBER 205, was used to perform the required calculations.

This analysis concludes that the preburner operational anomalies are not due to steady

state phenomena and must, therefore, be related to transient operational procedures.

-iv-
CI-FR-0084

INTRODUCTION

The fuel preburner of the space shuttle main engine (SSME), as shown schematically in

Fig. 1, consists of three major parts: the augumented spark igniter (ASI), injectors and

the combustion chamber. The fuel preburner itself is structurally supported by the hot

gas manifold to which it is welded and is a fuel cooled, double-walled chamber in which

hot gas (hydrogen-rich steam) is generated to power the high pressure turbopump. The

ASI initiates the combustion of the gaseous hydrogen fuel and liquid oxygen near the

injector elements on the faceplate. The fuel preburner operates at a high pressure of

5547.5 PSI and a temperature of 2000 °R after the gas/liquid propellants are mixed and

combusted. The preburned fuel then flows over a circular dome of the high-pressure fuel

turbopump (HPFTP) and into a flow passage of about one-third of the original cross

sectional area where it drives the turbines of the turbopump. These severe operating

conditions include the effects of incomplete mixing of the propellants and of possible

valve transient effects which can also produce local hot spots in the fuel preburner.

The concept of this modeling effort was to describe the bulk of the preburner flowfield.

Local hot spots at the stagnation point on the turbopump dome and on turbine blades may

be investigated with this modeL The cause of the hot spots on the dome of the

turbopump was suspected to be poor mixing of the oxygen injected in the ASI, i.e. a nearstoichiometric, hot streamline down the preburner centerline. Hot spots on turbine

blades are more difficult to explain, especially since some evidence of free oxygen has

also been observed. Modest radial and circumferential temperature variations are

expected at the turbine blades, but oxygen should be consumed under the very fuel-rich

flow conditions. Oxygen atomization and reaction should be rapid, and recent

Rocketdyne studies have confirmed that this is indeed the case (Ref. 2). Those studies

(Ref. 2) indicated that the oxygen was completely consumed in about one-third of a

baffle length in the streamwise direction. Instantaneous oxygen rich flows during start

up are believed by these authors to be the most likely source of turbine blade hot spots.

Based on these observations, a preburner model was developed.

The fuel preburner flowfield was calculated from a surface just downstream of the

injectors to the turbopump entrance. Groups of injectors were lumped to provide inflow

to computational node points so that multiple grid points per injector would not be

required. Upstream boundary conditions were developed such that flow into the boundary

-1-
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CI-FR-0084

nodes was represented with an orifice relation to relate hydrogen and oxygen manifold

pressure to local flow rates during steady operation. Delays in propellant vaporization

and combustion were investigated and found to be neglible. Mixing near the injectors

was described by the injector/lumping analysis; downstream mixing was by turbulent

transport only. The wall-cooled boundary layers were not modeled in this study. The

upstream boundary conditions were designed so that they could be used directly with an

unsteady analysis of the propellant manifolds. The numerical solution employed is also

unsteady, so the description of the manifold feed is all that is required to provide a

completely transient analysis. To investigate the fuel preburner flowfield, the

appropriate transport equations were numerically solved for both an axisymmetric grid

system and a three-dimensional system. Continuum's VAST (VAriational Solution of the

Transport equations) computer code was used to model the flowfield.

-3-
CI-FR-0084

VAST METHODOLOGY

Transport Equations

The VAST transport equations are written and solved in integral form (Ref. 3).

Conservation of Species

9

— f f - - I P. dV = - / (P. q + J.) ' dS ; i = l,...,m (1)

9t J l J l

v s

Where m is the number of species, P is the density, q is the velocity vector, J. is

the mass diffusive term, t is time and dS is the differential surface vector, while dV

is the differential volume.

Conservation of Momentum

3 i i s —

—3 f Pq dV = - / (Pqq - 3) • d§ ( 2 )

t L J. V "S

Where o" is the stress tensor.

Conservation of Energy

3

 r r - . _ - — I PE dV = - I (PEq * a • q + Q + B) • dS (3)

Where E is the intensive total energy, Q is the conductive heat flux vector and 8 is

the heat transferred by mass diffusion.

-4-
CI-FR-0084

The Numerical Analog

The VAST transport equations may be generalized as:

3

 f/ U.

at J. J

dV = - F

J '

j = 1,...,J (4)

Where J is the total number of transport equations and

U =

P.

PE

; F =

(P. q + Jj) ' dS

(Pq q - 5) ' dS

(PEq - 5 • q + Q + B) • dS

( 5 )

The integration domain may be subdivided into E finite elements and the VAST equations

reduced to :

(Au.Av/A t) e (6)

e=l

The flux Fj>e is evaluated over the surface of the elements and ( A U j A V / At ) e is the

total accumulation of the jth conserved quantity within element e. Since the sides of the

element terminate in nodes at which the discretized conserved quantity are known at a

previous time step, the new values of the conserved quantity at the next time step on the

nodes depend upon the allocation from the accumulations within the elements. The

conserved quantity at a node is therefore determined by assembling the contributions

from its surrounding elements. That is

Ay/At) n = .

k

e ?!

 C J,e, n F j , e

(7)

-5-
CI-FR-0084

Where n is the node number, k is the total number of elements surrounding the node and

£ . is an allocation parameter.

The assembled equation (7) is analogous to a finite difference expression in which the

spatial transformations are numerically imbedded in the analog. It may be interpreted as

a general form of the finite difference scheme since different finite difference schemes

can be derived with selected allocation parameters. Allocation parameters determine

the differencing scheme and the stability of the system.

VAST Differencing Scheme

The VAST differencing scheme is different from any existing schemes in that it does not

dictate a fixed allocation parameter throughout the course of integration, but changes

this parameter dynamically according to the variational principal of the real fluid

system. The original algorithm (Ref. 3) was based upon the satisfaction of the Second

Law of Thermodynamics. The advanced development (Ref. 4) requires that the rate of

entropy producton be maximized. It is a stronger argument than the Second Law, since if

the production rate of entropy is a maximum, it has to be greater than (or equal to) that

required by the Second Law. The VAST differencing scheme thus uses the transport

equations as equality constraints, subject to the objective of maximizing the entropy

production rate while dynamically determining the allocation parameters such that

maximum stability of the system can be achieved.

-6-
CI-FR-0084

TURBULENCE MODELING

Turbulence modeling involves the mathematical closure of the Reynolds stress terms

which result from time averaging the Navier-Stokes equations. Numerous turbulence

models have been developed in the recent years (Ref. 5). These models vary in

complexity and in their ability to accurately represent various flowfields. The

turbulence parameters in the models were determined empirically for the closure of one

or more systems, but are not necessarily applicable to other systems. Furthermore,

complicated turbulence modeling may introduce instability in numerical solutions. Since

the preburner flowfield is far more complicated than any other system for which

turbulence parameters have been determined, a straightforward and simple eddy

viscosity model with overall accuracy established by comparison to a single hydrogen/air

coaxial jet (Ref. 6) was used in the preburner calculations.

The effects of turbulence are modeled by an eddy viscosity coefficient

 U

T which enters

into the laminar Navier-Stokes equation by replacing molecular viscosity

 u

_ by

 u where

U = W + y (8)

The relationship between UT and y, is obtained from ducted turbulent flow

ln(uT/UL) = 0.89936 ln(Re) - 5.17515 (9)

where Re is the Reynolds number, based on the diameter of the inner jet.

The eddy thermal conductivity K and eddy mass diffusivity D are determined through the

input of the turbulent Prandtl number (Pr) and turbulent Schmidt Number (Sc).i.e.

Pr = - (10)

K

-7-
CI-FR-0084

and

Sc = (11)

D P

For a single coaxial jet, the potential core length of the inner jet, Xc, is determined

through Abramovich's empirical formula (Ref. 7). X approaches Xc when r^ approaches

zero.

X

= 1 - L- (12 )

and

L = ±

0.27(l-m)(0.21 4 + 0.144m)

(1+m)

(13)

where m is the ratio of the outer flowspeed to the inner flowspeed. In calculating L, a

minus sign is taken when m is greater than unity; r± is the radius of the inner tube and

r^p is the radius of the central potential core.

Turbulent eddy viscosities can be determined through equation (9) for the inner flow

Reynolds number Reinner and the outer annular flow Reynolds number Reouter. The

Reouter is based on hydraulic radius. The eddy viscosity model is then given by

Manner*

»»T(Reouter)-

l lT(Re inner)

In

y

T

(Reouter)

V^inner*

at XS

at X>X

(14)

-8-
CI-FR-0084

RESULTS OF THE SINGLE COAXIAL HYDROGEN/AIR JET CALCULATION

Axisymmetric calculations were conducted for the turbulent mixing of a parallel coaxial

hydrogen/air jet (Ref. 6). The inner hydrogen jet Mach number was 0.9 and the outer air

jet Mach number was 1.32. An initial velocity profile was given for the inlet condition.

A potential core length of 7.7 inner jet nozzle diameters was obtained through equation

12 and 13. Previous calculations (Ref. 8) used a grid system of 45 by 29 nodes and

reservoir upstream boundary conditions (RUBC). The results showed some oscillations in

the potential region for both the centerline velocity decay profile and the lateral

velocity profiles. These oscillations were postulated to be caused by the combination of

the coarse grid (9 longitudinal nodes in the potential region) and the RUBC.

A grid system of 53 by 30 nodes (15 longitudinal nodes in the potential region) with fixed

table value upstream boundary conditions (FTVUBC) was rerun for this report. Figure 2

shows the comparison of this new centerline velocity decay prediction against the axial

distance. Excellent matches of the predictions with the experimental data were

obtained. Previous centerline velocity decay is shown in Fig. 3 for comparison. Figure 4

shows the results of radial velocity and radial hydrogen mass fraction profiles at seven

different stations. The comparisons of the computational results with the experimental

data were quite good for a two "global" turbulence-level model. The good agreement in

radial profiles is not surprising since Chriss and Paulk (Ref. 9) have also obtained good

results with a constant radial eddy viscosity for these jets.

The improvement of this calculation over the previous one is believed to be due more to

the upstream boundary condition than the grid density. The inherent mass flow rate

fluctuation at the inlet caused by RUBC (Ref. 8) created oscillations in the flowfield and

also made it difficult for the calculation to converge.

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-15-
CI-FR-0084

COMBUSTION CHEMISTRY

Atmospheric pressure hydrogen/oxygen combustion kinetics were reviewed (Refs. 10, 11),

and the chemical kinetics rates suggested that the hydrogen/oxygen are expected to

combust very fast at higher pressures.

A one-step, reversible finite rate chemical kinetics model with Arrhenius rate expression

is developed in this study.

1

REACTION H + - O = H O (15)

222 2

The reaction rate of this model is verified in Fig. 5 and Fig. 6 for hydrogen/air ignition

delay time comparisons with a detailed twenty-eight reaction mechanism (Ref. 11). The

temperature histories of these two models agreed well at equivalence ratios of one and

two. These two figures also revealed that hydrogen/oxygen reaction is very fast and the

ignition delay times of hydrogen/air at rich combustion (<t > =2) is very close to those at

stoichmetric combustion.

This one-step, reversible finite rate chemical kinetics model was then used to predict the

ignition delay times for hydrogen/oxygen combustion at the fuel preburner operating

conditions (<f =7.4). The ignition delay time at fuel preburner operating pressure (377

atm), as shown in Fig. 7, is much shorter than that at atmospheric pressure; it is

practically instantaneous.

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-19-
CI-FR-0084

SUBSONIC BOUNDARY CONDITIONS

The fluid flow in the SSME fuel preburner is not only subsonic, but also has large

temperature gradients between the neighboring jets across the injector faceplate.

Previous studies (Ref. 8) have found that a fixed table value upstream boundary condition

(FTVUBC) in conjunction with a fixed integrated mass flow rate downstream boundary

condition is appropriate to describe the complicated subsonic preburner flowfield. The

fixed integrated mass flow rate downstream boundary condition regulates the steady

state mass flow rate throughout the fuel preburner flowfield, and the FTVUBC allows any

pressure disturbance originating from downstream to pass through the inlet boundary.

There is a loss of absolute conservation at the inlet and outlet boundary during the early

stage of integration which rapidly damps the transient phenomena and pressure

disturbances, but the flowfield resumes its absolute conservation as the computation

reaches steady state.

In this report, an injector simulation upstream boundary condition (ISUBC) is developed

to replace FTVUBC in order to better simulate the fuel preburner characteristics and to

anticipate the potential importance of manifold flow upstream of the injector

faceplate. The ISUBC has all the advantages of FTVUBC, yet further relates the

mainfold pressure to the preburner pressure, thus it is a tailor made upstream boundary

condition for the SSME fuel preburner simulation. The fixed integrated mass flow rate

downstream boundary condition was used in conjunction with ISUBC during this

computation. The methodology of the subsonic boundary conditions used in the fuel

preburner calculation is described as follows.

Fixed Integrated Mass Flow Rate Boundary Conditions

1. Let P ., P q, P E be free,

m

l

n J> 1 ?\

2. pq = pq , superscript n+1 stands for the new time step,

r PqdA

exi t

-20-
CI-FR-0084

3. pun + 1 = pqn + 1 COS 9

pvn + 1 = pqn + 1 SIN 9

where mj is the inlet mass flowrate and 9 means local flow angle.

This downstream boundary condition maintains a constant integrated mass flow rate

at the exit. A mass balance ratio is determined between the inlet mass flow rate and

the instantaneous exit integrated mass flow rate. This ratio is then applied to the

mass flux on the exit plane to keep a constant integrated mass flow rate throughout

the flow field.

Injector Simulation Upstream Boundary Condition

The SSME fuel preburner injector elements and manifold may be lumped together and

viewed as an orifice passage. The mathematical relationship between mass flow rate

and pressure drop between the two ends of the orifice can be derived by a

mechanical energy balance.

0.5

m = C [2P (P - P)] (16)

I o o

Where Cj is an effective discharge coefficient and subscript o denotes manifold

conditions. Cj can be estimated through steady state operating conditions or can be

obtained experimentally. The formal procedure is described as follows:

1. Let p., pq, PE be fre e

-21-
CI-FR-0084

2

1 Pq

2. P = ( Y- 1) (PE - )

2 P

n+1 0.5

3. pq = C [2p (P - P)] /S

l o o c

n + 1

4. p = p

i i ,T

n + 1

PE = PE

T

where y is the ratio of specific heats and the subscript represents table values

and Sc is the flow area.

5. Constant Flow Angle

Procedure 3 reveals that the ISUBC controls the flow rate at the injector face

via the pressure drop between manifold and injector faceplate. The flow rate

increases as the pressure at the injector face decreases and vice versa. The

irreversibility of ISUBC makes it absorb pressure disturbances sent from

downstream faster than the RUBC. The ISUBC regulates the mass flow rate

instead of resetting it, therefore causing less disturbances than FTVUBC.

The ISUBC also has the potential for separate fuel and oxidizer inlet flow rate

control, i.e., the injector manifold can be lumped into two or more orifices

instead of one. Separate discharge coefficients can be determined for each

orifice. ISUBC can also be applied to transient fuel preburner problems provided

that the pressure histories in the manifold are given.

-22-
CI-FR-0084

GENERAL FLOW FEATURES AND ASSUMPTIONS

The SSME fuel preburner consists of one ASI, 264 hydrogen/oxygen coaxial injector

elements and 588 tiny hydrogen bleed holes on the injector faceplate which feed

propellants into the combustion chamber. The major processes involved in this

complicated operation are atomization, vaporization, combustion and mixing. The fuel

preburner injector element is a baffled coaxial element that mixes the gaseous hydrogen

and liquid oxygen in correct proportions and uniformly distributes and injects the

propellants into the combustion chamber. The high velocity of the low density fuel,

relative to the density and velocity of the oxidizer, would produce a high rate of

atomization and thorough mixing (Ref. 1). It is then expected that the primary

atomization of liquid oxidizer through the injector elements would be completed in a

very short distance from the injector face. The operating variables of fuel preburner in

Table 1 and the initial properties in Table 2 indicate that the propellants are

"supercritical fluid" under these conditions. The atomized "supercritical fluid" drops

would gasify themselves without undergoing phase change (Ref. 12), i.e., instantaneous

vaporization or so called "shear stripping" (Ref. 13). The gasified and mixed propellants

would react almost instantaneously under fuel preburner operation conditions, which will

further help the atomization and mixing process due to the acceleration of the

combusted gases.

Due to the very rapid nature of the initial mixing, vaporization and combustion

processes, the atomization seems to be the rate controlling step which, in turn, could be

assumed very fast due to the high rate of atomization at the injector element. The best

engineering model to computationally represent the entire flowfield would be to model

the injector faceplate with concentric annual rings. Each ring represents a combustion

zone with fuel/oxidizer equivalence ratios determined by the layout of the injector

faceplate. The properties of the concentric ring jets will be calculated from equilibrium

calculations. The mixing of the ring jets will be simulated instead of resolving the

interaction of each individual jet of the injector element or holes. Such a model is

required because the resolution of the 264 injector elements and 588 hydrogen holes is

beyond present computational capability for practical applications.

Figure 8 shows the layout of the injector elements and holes on the injector faceplate.

Four major combustion zones can be grouped together according to the sizes and

-23-
CI-FR-0084

locations of the injector elements and holes. They are igniter zone, AB zone, CD zone

and EFGH zone. The mass flow rate is assumed to be evenly distributed among all the

injector elements and holes. There is an additional baffle flow zone the in threedimensional modeL

-24-
CI-FR-0084

Table 1: Operating variables for fuel preburner at FPL of 1.09

Pressure (PSIA) Temperature (°R)

Manifold Hydrogen Inlet 6023 278

Mainfold Oxygen Inlet 7135 213

Igniter Hydrogen Inlet 6781 102

Igniter Oxygen Inlet 7212 213

FPB Chamber 5547 1990

Table 2: Critical properties for hydrogen, oxygen and water

Pressure (PSIA) Temperature (°R)

Hydrogen 188 60

Oxygen 732 278

Water 3199 1165

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-26-
CI-FR-0084

AXISYMMETRIC MODELING OF THE FUEL PREBURNER

Grid Definition

Two grid systems were used in the axisymmetric fuel preburner calculations. The coarse

grid definition, as shown in Fig. 9 was composed of two regions. Region I contains the

igniter flow region and main combustion chamber between the injector face and 5.55 in.

below the injector face. Region n covers the rest of the chamber. There are 540 nodes

in Region I; 126 of them are null nodes. Region n has 350 nodes. The coarse grid

definition therefore has 890 nodes in total; 764 of them are active nodes. The final grid

definition, as shown in Fig. 10, is composed of three regions. Region I is the igniter flow

region. Region n contains the combustion chamber between injector face and 5.55 in.

below the injector faceplate. Region III covers the rest of the combustion chamber.

There are 5,965 nodes in the fine grid definition. Region I has 160 nodes, region n has

1,800 nodes and Region II has 4,005 nodes. No null nodes were used for the fine grid

definition.

-27-
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-28-
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-29-
CI-FR-0084

Boundary and Inlet Conditions

Adiabatic, free-slip conditions were assigned to all the combustion chamber walls on the

assumption that the turbulent wall boundary layer is very thin and there is no heat

transfer to the environment. Heat exchange across the preburner liner can be modeled,

but this effect was not included in this study. One null region was used in the coarse grid

definition. Table 3 shows the inlet and boundary conditions used for the axisymmetric

modeL Since the mixing near the injectors was described by the injector/lumping

analysis a turbulent viscosity of 4.79 E-3 Ib/ft-sec was obtained from equation 9 and was

used throughout the combustion chamber. The same value of viscosity was also used in

the three-dimensional analysis. The fuel/oxidizer equivalence ratios for the igniter flow

zone, AB zone, CD zone and EFGH zone are 18, 9.5, 7.8 and 6.9 respectively. The

injector simulation upstream boundary conditions were used for all four inlet zones. A

fixed integrated mass flow boundary condition (FIMFBC) was used for the exit boundary.

-30-
Table 3 Inlet and Boundary Conditions for Axisymmetric Model

CI-FR-0084

Zone

•

m (Ib/sec)

aH2

aH20

T(°R)

P(lbf/ft2)

P0(lbf/ft2)

C

I

q (ft/sec)

B.C.

Igniter Flow

1.55

0.65853

0.34147

1112.4

798804

1007452

3.25E-4

191

ISUBC

AB

22.43

0.48678

0.51322

1796.4

798804

947357

6.45E-3

257

ISUBC

CD

36.53

0.43272

0.56728

2098.8

798804

947357

1.08E-2

282

ISUBC

EFGH

114.86

0.39854

0.60141

2298.6

798804

947357

3.45E-2

333

ISUBC

Exit

175.38

FIMRBC

-31-
CI-FR-0084

Results for the Axisymmetric Fuel Preburner Models

Figure 11 shows the flow vectors for the coarse grid calculation. There is a flow stagnation

region formed near the top of the fuel turbopump dome. The four streams of combusted

products flowed down the combustion chamber and gradually merged with each other as the

flow passage narrowed over the fuel turbopump' dome. The flow speed of each streamline

reached its maximum at the exit. The mass flow rate at each cross sectional plane of the fuel

preburner was constant at steady state due to the fixed integrated mass flow rate boundary

condition at the exit. The flow vectors of the fine grid calculation are similar to that of the

coarse grid calculation.

The flow area contraction ratio of the main chamber to the exit is about three to one. The

large flow area contraction ratio is reported to cause a high level of secondary recirculation

(Ref. 14) and, therefore, has the tendency to smooth out temperature profile variations. Figure

12 and Fig. 13 show the computational results for the temperature profile and hydrogen mass

fraction profile of the coarse grid calculation. The results seemed to support the expectation

that the large flow area contraction ratio does indeed tend to cause the four streams of

combustion flows to mix completely at the exit section of the fuel preburner.

The computational results of the fine grid calculation show the flow to be mainly streamline.

The temperature profile in Fig. 14 showed that although a certain amount of mixing occurred,

temperature variations did exist at the exit of the preburner. Streamlines of colder

temperatures climbed over the fuel turbopump dome to the exit and so did the hot, outer

combustion streamlines. An averaged, radial temperature profile at the exit (Fig. 15) showed a

300 °R variation. Figure 16 shows the result of the hydrogen mass fraction profile of the fine

grid calculation. The mass fraction streamlines showed hydrogen mass fraction variations at

the fuel preburner exit. The discrepancies between the results of the coarse grid calculation

and those of the fine grid calculation show the effect of numerical diffusion due to

discretization. The results of the fine grid calculation should be used as the input for turbine

flowfield calculation. About 1000 CPU seconds were used for coarse grid case and 8000 CPU

seconds were used for fine grid case on CYBER 205 to reach steady state.

-32-
ORIGINAL PAGE IS

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-38-
CI-FR-0084

THREE-DIMENSIONAL MODELING OF THE FUEL PREBDRNER

Grid Definition

Three grid regions were used to construct the grid system for three-dimensional fuel

preburner modeling. A front view of the three-dimensional fuel preburner grid with a 30

deg. angle is shown in Fig. 17. Region I is the augmented spark igniter chamber; Region

II comprises the major part of the fuel preburner combustion chamber, the injector

faceplate and three flow-straightening cooling-baffles; Region ffl represents the

flow field for the combusted gases that flows around the hemispheric high pressure fuel

turbopump dome and eventually leaves the fuel preburner exit to drive the turbines. The

coordinate system is such that the n 3 coordinate coincides with the central axis of the

fuel preburner, n coordinate points radially from center to the wall, and n«

coordinate goes circumferentially around the centerline. This way the 3-D grid system

can be closed by starting with a slice of the 2-D grid system and move it

axisymmetrically around the centerline for 360 deg. to overlap with the original 2-D

slice. Region n and Region III have to be separated in order to get the exact sizes and

shapes for the baffles and the hemispheric dome. Region I was separated from Region n

to save a large number of nodes. Notice that there are no nodal points inside the

turbopump dome. The total number of nodes used in the computation was 21,638.

Figure 18 shows a top view of the 3-D fuel preburner grid system. Baffle 1 is located at

330 deg. baffle 2 is located at 90 deg., and baffle 3 is located at 210 deg.

-39-
ORIGINAL PAGE IS

OF POOR QUALfTY

Fig. 17 Front View of the 3-D SSME Fuel Preburner

Grid Definition With a 30 Degree Angle

-40-
BAFFLE 2

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Fig. 18 Top View of the 3-D SSME Fuel Preburne r Grid Definition

-41-
CI-FR-0084

Boundary and Inlet Conditions

Adiabatic free slip conditions were assigned to all the combustion chamber walls for the

purposes of this study. Null regions were assigned to the inside of the three baffles.

Table 4 shows the inlet and boundary conditions for the three-dimensional model. The

injector simulation upstream boundary condition was used on all four inlet combustion

zones on the injector faceplate. A fixed inlet boundary condition was used for coolant

hydrogen flow on the baffle. Fixed mass flow rate boundary conditions were applied to

the exit. Swirling components were added to the ASI inlet velocity vectors to simulate

hydrogen in flow.

-42-
CI-FR-0084

Table 4 Inlet and Boundary Conditions for Three-Dimensional Model

Zone Baffle Flow Igniter Flow AB CD EFGH Exit

m (Ib/sec)

«H2

°

H

2°

T(°R)

P(lbf/ft2)

P0(lbf/ft2)

q(ft/sec)

B.C.

3.15

1.0

0.0

300

798804

59

FIBC

1.55

0.65853

0.34147

1112.4

798804

1007452

3.25E-4

191

ISUBC

20.98

0.45129

0.54871

1985.4

798804

947357

6.16E-3

398

ISUBC

35.97

0.42382

0.57618

2143.8

798804

947357

1.07E-2

284

ISUBC

113.73

0.39259

0.60741

2338.2

798804

947357

3.43E-2

417

ISUBC

175.38

FIMRBC

-43-
CI-FR-0084

Results for the Three-Dimensional Fuel Preburner Model

The three-dimensional model provides more insight into the fuel preburner operation than

the axisymmetric model in that swirling ASI flow effects, baffle flow straightening

effects, baffle coolant flow effects and asymmetric flow effects due to unbalanced

heating or cooling can be studied. Figure 19 shows the swirling components of the ASI

inlet flow. The swirling comes from the eight tangential fuel injectors around the center

oxidizer jets. The rich swirling ASI flow not only ignites the main injector flow but also

protects the ASI combustion chamber walls. Figure 20 shows a three-dimensional plot

view of the flow vectors. Seven slices of the grid system constitute the plot view with a

30 deg. front view angle. The plot view consists of a section of the ASI inlet, a section

of the injector faceplate, the left section of the outside wall, an internal section which

contains the baffle, another internal piece which sits between two baffles, a section of

the surface of the turbopump and the exit. The rectangular area in one of the internal

sections is one of the baffles. There is a stagnated flow region near the top of the

turbopump dome. The coolant hydrogen flow comes down the baffle slowly and mixes

gradually with bulk flow while accelerating toward the exit. The slice of the internal

flow located between two baffles looks similar to that of an axisymmetric calculation.

The swirling component of ASI flow loses its momentum below the injector faceplate,

partially due to the baffle flow straightening effect, and partially due to its small flow

rate.

Figure 21 shows the temperature contours for the same plot view. The cooling effect of

the baffle coolant flow can be seen clearly from the internal section on the right. The

1850 °R isotherm traced on the surface of the turbopump dome shows a combined

influence from the igniter flow and the coolant baffle flow. The lack of variation in

temperature gradient at the exit indicates the existence of a coarse grid smearing

effect. Figure 22 shows a cross-sectional view of temperature contours at the bottom of

the baffles. The residual effect of the swirling can be seen close to the baffles. Figure

23 shows another cross-sectional view 1.63 in. above the top of the turbopump. The

cooler temperature indicates the influence of the baffle coolant flow.

Figure 24 compares the predicted circumferential temperature profile near the preburner

exit with that from hot fire test data (Ref. 14). The hot fire test was conducted with

SSME component preburner assemblies. The preburner was assembled with a nozzle

assembly which acted as a turbine simulator. The thermocouple ports in the test were

-44-
CI-FR-0084

located on two planes in the turbine inlet zone. Twenty-four temperature sensing

positions were located in the lower plane which was approximately 10.75 in. downstream

from the injector face. Half of them were installed with sensing tips extending into the

discharge nozzles, and the temperatures obtained are called "nozzle" temperatures. The

remaining half were located midway between discharge nozzle and are called "standard"

temperatures. Sixteen additional thermocouple ports were equally spaced around the

nozzle assembly on a plane about 9.5 in. from the injector face. These are referred to as

"special" temperatures. The predicted temperature profile is an average of the

temperatures of the nodes in between these two planes. The predicted temperature

profile showed similar trends to those of test data. The temperatures are near their

minimum under the baffles, and are near their maximum in between baffles. The

predicted temperatures are somewhat higher than those of the test data. This is because

the hot fire test was operated at an earlier version of the design which has a higher

fuel/oxidizer equivalence ratio of 8.96 rather than the 7.65 used by the present

computation. Nevertheless, the agreement of the trend and the closeness of the data

level verify the validity of the model. This run takes about 29,000 CPU seconds on the

CYBER 205 to reach steady state.

-45-
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Fig. 21 3-D SSME Fuel Preburner Temperature

-48-
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-51-
CI-FR-0084

CONCLUSIONS

1. The result of a recent paper (Ref. 2), concerning the analysis of a single SSME fuel

preburner injector element, has confirmed our model assumptions about fast

atomization, immediate vaporization and instantaneous combustion in the fuel

preburner. They found a short liquid jet length of about 2 cm, and the existence of

few droplets beyond the tip of the liquid jet. The length of 2 cm amounts to just 6

percent of the total length of the fuel preburner chamber. The fact that the curves

for percentage oxidizer atomization, vaporization and reaction reached 100 percent

at 2 cm below the injection inlet and are practically identical, points out the

instantaneous nature of vaporization and combustion during preburner operation.

2. The agreement in trend and data level of the circumferential exit temperature

between the three-dimensional model and hot fire test data supports the accuracy of

the computational model.

3. Due to the high rate of atomization, immediate vaporization and instantaneous

combustion, no oxygen can survive the length of the combustion chamber to reach

the Kaiser Hat on the top of the turbopump dome, or move into the convergent

section and escape to burn turbine blades. The steady state flowfield of the

preburner should be as our model suggests: striated combustion flows come down

from the injector faceplate with hotter temperatures toward the outside wall due to

the layout of the injectors, is cooled by coolant hydrogen from three equally spaced

baffles, converges as it enters the convergent section and finally leaves for the

turbines. The rich igniter flow produces colder flow, which not only protects the

igniter chamber wall, but also cools down the turbopump dome surface. Therefore,

there are no apparent hot spots formed during the steady state operation, although

the axisymmetric model predicted temperature variation on the exit. The model

assumes instantaneous mixing of the near injector region, and the turbulent

calculation indicates very little additional mixing as the flow moves downstream.

This important observation should be experimentally verified.

4. Local hot spots on the turbine blades are postulated to come from transient

operation, i.e. flow leads and lags may cause near stoichiometric mixtures

downstream of the preburner during start-up of the engine.

-52-
CI-FR-0084

5. The SSME fuel preburner flow field was successfully modeled by axisym metric and

three-dimensional models. The model generates realistic flows that can be used to

understand the complicated preburner flowfield and aid the assessment of the design

modifications.

-53-
CI-FR-0084

RECOMMENDED MEASUREMENTS STUDY

As a result of flowfields described by the analyses reported herein, the following

measurements are recommended:

1. The large temperature and mixture ratio gradients which are predicted at steadystate should be verified by direct measurements. Since it is very difficult to insert

adequate probes for such measurements, optical emission/absorption measurements

using fiber optics as radiation paths should be considered. The Test Bed Engine

program may be used to accomplish these measurements.

2. Transient preburner exit temperatures should be measured in the test bed engine

program.

3. The large differences seen in near-injector mixing and in highly turbulent parallel

stream mixing should be investigated experimentally in a geometrically simple

combustor. A laboratory combustor flow in which a few, easily varied injector

elements can be used should be studied. Transparent side walls should also be used.

The turbulent mixing and reaction progress must be quantitatively measured both

optically and with probes.

-54-
CI-FR-0084

REFERENCES

1. Space Transportation System Technical Manual - SSME Description and

Operation, E41000/RSS-8559-1-1-1, Rocketdyne Division, Rockwell

International Corporation, 5 April, 1982.

2. Liang, P.Y., R.J. Jensen and Y.M. Chang, "Numerical Analysis of SSME Fuel

Preburner Injector Atomization and Combustion Processes,11 paper presented

at AIAA conference, Reno, Nevada, January 1986.

3. (a) Prozan, R.J., "A Variational Principle for Compressible Fluid Mechanics:

Discussion of One-Dimensional Theory," NASA CR3526 1982.

(b) Prozan, R.J., "A Variational Principle for Compressible Fluid Mechanics:

Discussion of the Multi-Dimensional Theory", NASA CR3614, 1982.

4. Prozan, R.J., "Hypothesis of a Variational Principle for Compressible FLuid

Mechanics", CI-TR-0086, 1985.

5. Sindir, M.M., and P.T. Harsha, "Assessment of Turbulence Models for Scram jet

Flowfields", NASA CR3643, 1982.

6. Eggers, J.M., "Turbulent Mixing of Coaxial Compressible Hydrogen-Air Jets",

NASA TN D-6487, 1971.

7. Abramovich, G.N., The Theory of Turbulent Jets, MIT Press, 1963.

8. Wang, T.S. and R.C. Farmer, "Computational Analysis of the SSME Fuel

Preburner Flow," Interim report, CI-IR-0078, January 21, 1985.

9. Chriss, D. and R.A. Paulk, "An Experimental Investigation of Subsonic Coaxial

Free Turbulent Mixing", AFOSR-TR-72-0237, 1972.

10. Edelman, R.B., R.C. Farmer, and T.S. Wang, "Combustion and Emissions of

Synthetic Fuel Components, Analysis and Modeling", Combustion of Synthetic

Fuels (W. Bartok, ed.) ACS Symposium Series 217, American Chemical

Society, pp. 29-48, 1983.

11. Rogers, R.C. and W. Chinitz, "Using a Global Hydrogen-Air Combustion

Model in Turbulent Reacting Flow Calculations", AIAA Journal, VoL 21, No. 4,

pp.586-592, 1983.

12. Smith, J.M. and H.C. Van Ness, Introduction to Chemical Engineering

Thermodynamics, 3rd edition, McGraw Hill, Inc., 1975.

13. Rabin, E. and R.8. Lawhead, "The Motion and Shattering of Burning and NonBurning Propellent Drops," AFOSR-TN-59-129, 1959.

14. Fleck, D.L., "Design Verification Complete Package—Hot Gas Temperature,"

RSS-305-22, October 23, 1979.

-55-