M.E.M. Stewart and A. Suresh

QSS Group, Inc., Cleveland, Ohio

M.S. Liou and A.K. Owen

Glenn Research Center, Cleveland, Ohio

D.G. Messitt

Aerojet, Sacramento, California

Multidisciplinary Analysis of a

Hypersonic Engine

NASA/TM—2002-211971

October 2002

AIAA–2002–5127
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M.E.M. Stewart and A. Suresh

QSS Group, Inc., Cleveland, Ohio

M.S. Liou and A.K. Owen

Glenn Research Center, Cleveland, Ohio

D.G. Messitt

Aerojet, Sacramento, California

Multidisciplinary Analysis of a

Hypersonic Engine

NASA/TM—2002-211971

October 2002

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

11th International Conference on Space Planes and Hypersonic Systems and Technologies

cosponsored by the American Institute of Aeronautics and Astronautics and

the Association Aeronautique et Astronautique de France

Orleans, France, September 29–October 4, 2002

AIAA–2002–5127
Acknowledgments

This work was supported by the NASA Computing and Information Communications Technology (CICT) program

through the Computing and Interdisciplinary Systems Office (CISO) (contract NAS3–00145) at NASA Glenn

Research Center. The authors would like to thank Jeffrey A. White, NASA Langley Research Center, for help

 with VULCAN.

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Available electronically at http://gltrs.grc.nasa.gov
NASA/TMó2002-211971 1 

MULTIDISCIPLINARY ANALYSIS OF A HYPERSONIC ENGINE

M.E.M. Stewart and A. Suresh 

QSS Group, Inc. 

Cleveland, Ohio 44135 

Mark.E.Stewart@grc.nasa.gov

M.S. Liou and A.K. Owen 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

D.G. Messitt 

Aerojet 

Sacramento, California 95813 

Abstract 

This paper describes implementation of a technique 

used to obtain a high fidelity fluid-thermal-structural 

solution of a combined cycle engine at its scram design 

point. Single-discipline simulations are insufficient 

here since interactions from other disciplines are 

significant. Using off-the-shelf, validated solvers for 

the fluid, chemistry, thermal, and structural solutions, 

this approach couples together their results to obtain 

consistent solutions. 

Introduction 

To reduce the cost of access to space, NASA has 

focused on several propulsion concepts. In one longerterm view, the ISTAR program is examining the 

Rocket Based Combined Cycle [1,2,3] (RBCC) 

concept that integrates a rocket and a ram/scram jet. In 

particular, the low speed and ex-atmospheric 

advantages of rocket propulsion are combined with the 

higher specific impulse of air breathing propulsion to 

obtain a more efficient propulsion system. A concept 

design analyzed here is a strutjet engine that alternates 

rocket containing struts with combustor ducts as 

indicated in Figure 1. 

Particularly with high fidelity analysis, a common 

design practice is to neglect some multidisciplinary 

interactions. While this practice is adequate in many 

instances, a successful RBCC design is a careful 

balance between aerodynamics, combustion, thermal 

management, structural and weight requirements. 

Consistent multidisciplinary solutions should capture 

these interactions, identify their consequences, and 

consequently play a role in design. 

The techniques implemented here apply not only to 

RBCC designs but to a range of problems where 

multidisciplinary interactions are significant. In 

particular, these techniques may be readily applied to 

TBCC designs. A long term goal of this work is 

developing a toolkit that simplifies the 

multidisciplinary coupling of off-the-shelf codes. 

The following sections present the three component 

simulations, details of the coupling of these simulations, important coupling issues, results, and a discussion of the added cost of multidisciplinary analysis. 

Component Simulations 

The present work involves three steady, threedimensional simulations for ISTAR engine 

components: a fluid simulation of the approach flow 

over the vehicle forebody and engine duct inlet, a 

fluid-chemistry simulation for the combustor, and a 

thermal-structural simulation of the engine walls. 

Vehicle Forebody and Engine Inlet:

The hypersonic approach flow over the vehicle 

forebody and into the engine inlet is calculated at the 

scram design point with the Navier-Stokes code, 

OVERFLOW [4]. A κ-ω turbulence model is used with 

boundary layer grid resolution of y+

~1 (at the first 

node off the wall) [4,5]. OVERFLOW modifications 

allow simulation of equilibrium chemistry air. All 

vehicle surfaces that can influence the engine inlet 

airflow are included in a Chimera (overlapping) 
NASA/TMó2002-211971 2 

Figure 1: ISTAR Vehicle concept and Strutjet Engine [1] (side and front views) 

Figure 2: Approach flow Mach contours along the symmetry surfaces of the vehicle forebody and one engine strut-to-strut 

flowpath.

structured grid system (5 blocks; 9×105

 cells) 

generated with GRIDGEN [6]. Here, Chimera grids 

simplify complex geometry grid generation, and 

accommodate some design changes. Interpolation 

quantities for this Chimera grid system are calculated 

with PEGASUS [7]. Centerline symmetry is assumed 

for both the vehicle and the single strut-to-strut flow 

path simulated. Wall temperatures must be 

specifiedinitially a guess and on subsequent cycles 

temperatures are interpolated from the solidís thermalstructural model. The forebody/inlet configuration and 

an approach flow solution are shown in Figure 2. 

Combustor Fluid-Chemistry Analysis:

Scram combustion within the engine is simulated using 

the Navier-Stokes plus finite rate chemistry code, 

VULCAN [8]. A κ-ω turbulence model with wallfunctions is used with grid resolution of the boundary

layer to y+

max<500. Radiation effects are not included. 

The combustor inflow is supersonic and calculated 

from the approach solution by interpolating solution 

quantities between grids. 

Within the engine, the cascade fuel injectors are modeled as single triangular slots with area, mass flow, and 

momentum equal to the actual injectors. Combustion is 

simulated with a 6-species, 3-step finite-rate Ethylene 

model. Although it was not part of the preliminary 

CAD model, a flame holding cavity was added to 

facilitate and sustain combustion. Wall temperatures 

must be specifiedóthe initial value is a guess, and on 

subsequent cycles temperatures are interpolated from 

the thermal model. The grid contains five composite 

(non-overlapping) blocks (1.9×105

 cells), and 

centerline symmetry is assumed. A typical combustor 

configuration and a solution are shown in Figure 3.
NASA/TMó2002-211971 3 

Figure 3: Combustor duct cutaway showing fuel mass fraction iso-surfaces colored by temperature. Centerline symmetry is 

assumed. 

Engine Strut Thermal and Structural Analyses:

Both the thermal and structural simulations are 

performed individually using ANSYS [9], a 

commercially available finite element solver. From 

engine geometry CAD files, a three-dimensional, 

unstructured, tetrahedral thermal-structural mesh was 

created within ANSYS (1.3×105

 nodes, 8.6×104

tetrahedra); shell elements, although simpler, cannot 

capture the normal heat conduction that is of interest 

here. Centerline symmetry was assumed. 

Temperature-dependent material properties for Inconel 

625 and Titanium β21S are taken from manufacturerís 

specifications [10,11]. A thermal barrier coating on the 

engine strut surface is modeled with homogenized 

material properties. Coolant passages are modeled by 

dividing the geometry into two-layers at coolant 

surfaces; the coolant temperature is applied at this 

bilayer surface (Figure 4). This model neglects details 

of heat conduction around each coolant passage. 

Surface heat fluxes and pressures are calculated 

from the fluid solutions, interpolated onto this thermalFigure 4: Coolant Passages (left) are modeled with a bilayer 

material (right); the coolant temperature is applied on the 

intermediate surface.

structural model, and used as boundary conditions in 

each ANSYS analysis. The coolant temperature 

boundary condition is specified in the thermal analysis 

as either a fixed temperature, or as a temperature 

distribution calculated after integrating the heat flux 

along each coolant channel. After the thermal analysis 

of this model, the resulting wall temperatures are 

interpolated onto the surfaces of the fluid grids and 

used as boundary conditions in the fluid calculations. 

Similarly, surface deflections from the structural 

analysis can be interpolated and used to deform the 

fluid grids. A typical thermal solution is shown in 

Figure 5, and a typical structural solution is shown in 

Figure 6. 
NASA/TMó2002-211971 4 

Figure 5: Engine strut thermal model showing temperature contours for scram combustion. The five fuel-injectors are the small,

dark triangular regions. 

Figure 6: Engine strut structural model showing deflection contours at the scram design point. Deflections are exaggerated. 
NASA/TMó2002-211971 5 

Coupling Procedures 

The objective of coupling these three simulations 

together is to ensure a consistent solution for the 

engine: flow quantities are the same where the fluid 

codes meet, heat fluxes and temperatures are the same 

where fluid and thermal codes meet, and the deflected 

walls are the same as the fluid boundaries. Three 

coupling procedures are necessary here; between the 

inlet and combustor calculations there is a fluid-fluid 

coupling, and between the flowpath and the solid walls 

both a fluid-thermal and a fluid-structural coupling 

exist. 

Fluid-Fluid Coupling

The inlet (OVERFLOW) and combustor (VULCAN) 

fluid calculations overlap near the throat, and coupling 

ensures that flow quantities are consistent there. Since 

the flow is supersonic and boundary layers are 

attached, the downstream influence is assumed to be 

negligible; outflow values of inlet solution variables 

are interpolated onto the combustor grid. 

 

There are three interpolation challenges in this 

coupling. First, different codes may use different flow 

or turbulence variables, nondimensionalizations, 

and/or units; transformation of variables may be 

necessary. Here, κ-ω turbulence models were chosen 

in both calculations. However, for a faster turnaround 

time, wall-functions were chosen in VULCAN; 

OVERFLOW has only a low-Reynolds number 

(integration to wall) turbulence model. Second, the 

interpolation of ω was complicated by its singular 

behavior in boundary layers. Interpolation of the 

turbulent viscosity, ρκ/ω, was better behaved. Third, 

interpolation in the highly resolved boundary layer 

meshes near curved walls can be difficult. In Figure 7, 

discretely sampled points on one mesh (dashed line) 

can lie outside the other mesh (solid line); 

interpolation is not possible, and extrapolation from 

nearest neighbors is required. 

Figure 7: Interpolation between two discrete grids (solid and 

dashed lines) fails at point A on a curved boundary (dotted 

line). Extrapolation from nearest neighbors is required. 

Fluid-Thermal Coupling

The correct thermal boundary condition is continuous 

heat fluxes and temperatures at the interface between 

fluid and thermal codes; however, calculating heat 

fluxes from a fluid solution and applying them in a 

thermal calculation, will not satisfy this condition, in 

general. 

To obtain consistent thermal solutions, the current 

approach is to iterate between the fluid and thermal 

solvers [12,13,14]. Starting with a guess temperature 

for the fluid-solid interface, a fluid solution is 

obtained, interface heat fluxes are calculated and 

imposed on the thermal solver. The resulting thermal 

solution revises the interface temperature. This cycle 

continues to convergence, and in practical problems, 

this procedure converges within 10 iterations [12]. 

There are three challenges in this coupling. First, the 

convergence of this procedure is sensitive to several 

factors including material properties and the initial 

guess wall temperature. Some work has not required 

under-relaxation for convergence [12], while other 

researchers have used it [14,15]. In the current work, 

interface temperatures oscillated during the fluidthermal iteration both for VULCAN and 

OVERFLOW. Under-relaxation (w=0.25−0.5) of the 

interface temperatures calculated by ANSYS allowed 

convergence. The appendix presents a theoretical 

analysis that predicts oscillations, sensitivity to the 

material thermal conductivities, and the need for 

under-relaxation to improve convergence. 

Convergence was also improved by using VULCANís 

thermally mixed boundary condition to generate the 

initial guess temperature distribution. This boundary 

condition couples heat fluxes at the wall with a onedimensional solid wall heat conduction analysis [15]. 

Second, the calculation of accurate heat fluxes is 

challenging [13] for OVERFLOW. The heat flux 

calculation involves a difference of flow variables 

(which decreases the order-of-accuracy) performed in 

the highly refined boundary layer grid. Any lack of 

smoothness in the grid contributes to noisy fluxes; the 

grid singular line in the rounded strut corner is a 

source of noise. 

Third, the turbulence model has an influence on the 

wall heat fluxes. VULCAN uses wall-functions and a 

coarser boundary layer grid resolution. Here, the heat 

flux is calculated from a functional representation of 

A 
NASA/TMó2002-211971 6 

the boundary layer profile. In practice, VULCAN heat 

fluxes are less noisy. However, a debate exists about 

the relative accuracy of heat fluxes calculated from 

wall-function and low-Reynolds number turbulence 

models. Even in a geometrically simple test problem 

(Mach 3 flow across a constant temperature flat plate), 

heat flux predictions from VULCAN and 

OVERFLOW differ by 11 to 15 percent. Further, in 

the engine duct where the converged OVERFLOW 

and VULCAN fluid-thermal simulations overlap, the 

heat fluxes have disparate values. 

Fluid-Structural Coupling

Like the fluid-thermal coupling, the correct fluidstructural interface condition is consistent pressures 

and deflections between fluid and structural codes. 

The common practice of neglecting surface deflections 

in the fluid simulations does not satisfy this condition, 

in general. 

 

To find consistent structural solutions, pressures from 

the converged fluid-thermal simulations are 

interpolated onto the fluid-solid interface of the 

structural grid; these pressures are boundary 

conditions in the structural analysis. From the thermal 

analysis, the temperature distribution in the solid is 

used in the structural analysis. Deflections calculated 

in the structural analysis are interpolated back to the 

fluid grids and used to deform these grids. This cycle 

is continued to convergence. 

There are two interpolation challenges in this 

coupling. First, the engine ramp and engine strut 

deflect independently, since they are not attached 

(Figure 8). Deformations are discontinuous across the 

wall-ramp gap, and the interpolation procedure must 

not confuse points on different components. 

Consequently, when searching for an interpolation 

stencil, a restriction is required: interpolation must be 

to the same engine component. 

Figure 8: The engine ramp and strut deflect independently; 

interpolation must be to the same engine component. 

Second, deforming fluid grids becomes difficult as 

deformationsóparticularly shear (tangent to wall) 

deformationsóexceed the grid spacing. Since 

VULCAN uses wall functions, wall spacing is coarse 

and grid deformation is straightforward. OVERFLOW 

grids have finer wall spacing and must also be of 

sufficient quality for CHIMERA/PEGSUS 

interpolation. These two constraints present a 

challenge. 

Results 

Fluidóincluding combustionó(Figures 2, 3), thermal 

(Figure 5), and structural solutions (Figure 6) have 

been obtained for this ISTAR engine configuration at 

its scram design point. The fluid and thermal solutions 

have been converged so that temperatures and heat 

fluxes are consistent at the fluid-solid interface. The 

fluid and structural simulations have been iterated 

through one cycle. 

The fluid-thermal iteration substantially changes the 

engine strut wall temperatures. The initial, uniform, 

wall temperature of the strut was taken to be 1560 °R 

for both OVERFLOW and VULCAN. The L2 norm, 

(∑ (∆Τ)

2

/N)1/2, of the temperature change, ∆Τ, between 

the initial and converged temperatures was 500 °R. 

The heat fluxes calculated before and after the fluidthermal convergence are substantially different. As 

expected the qualitative details of the duct flow 

changed slightly. 

These computational results reveal quantitative details 

of inlet performance, engine combustion, heat transferthermal management, and structural deflections and 

stresses. If these analyses were performed 

concurrently with the early design process, these 

results would have been valuable. These techniques 

compliment cycle analyses, and one- and twodimensional computations, and help understand wind 

tunnel data and flight engine design. Also, these 

results can provide insights into design trade-offs. As 

with any CFD calculation, one must consider the 

limitations of the numerical methods, computational 

grid, and physical models. 

The Cost of Multidisciplinary Analysis 

The manual effort to setup and perfect single 

discipline simulations is substantialon the order of 

months. This effort is dominated by the manual effort 

of structured, fluid grid generation. However, it is 

Ramp

Duct Wall Grid 
NASA/TMó2002-211971 7 

important to distinguish the added cost of performing a 

multidisciplinary analysisówith these single 

discipline simulations in hand. This cost can be 

broken into two components: the additional 

computation time, and the setup time for the iterative 

coupling. While computation time approximately 

doubles, the increase in setup time is hard to quantify 

and may be reduced by a coupling toolkit. 

Estimating the added computation time for a 

multidisciplinary simulation is also complicated by the 

disparate execution times of the component solvers. 

The thermal-structural model and interpolation require 

only minutes; while the fluid-combustion models 

require on the order of tens of hours. Some 

improvements may be possible. More aggressive 

parallel execution of the fluid solvers may reduce 

turnaround time. Further, in this preliminary work the 

fluid codes were completely converged within each 

fluid-thermal iteration; with such disparate time scales, 

this restriction may waste computational resources, 

especially in the initial iterations when large changes 

occur on the boundaries. 

Setting up these multidisciplinary couplings involves 

working with file formats to output and input solution 

variables; transforming variables to correct for 

different units, nondimensionalizations, and even 

different solution variables; massaging files to identify 

and format interpolation surfaces; interpolating these 

variables between grids; and carefully overseeing the 

entire process. A goal of this work is automating these 

steps, where possible, and creating a general toolkit for 

coupling off-the-shelf simulation codes. 

Implementation of this coupling procedure has 

revealed obstacles to achieving this goal. First, 

interpolation is challenging particularly in boundary 

layer meshes, at boundaries, and where components 

meet. Second, variable transformations are often 

unavoidable; different codes use different units and 

nondimensionalization of solution variables, (i.e. 

different turbulence models). Third, the calculation of 

heat fluxes is problematic. Fourth, when shear 

deformations are larger than grid spacing, deforming 

grids can be challenging. 

Summary 

This paper describes a procedure for obtaining 

consistent fluid-thermal-structural solutions. The 

procedure is intended to handle off-the-shelf, singlediscipline solvers with limited or no access to solver 

source code. The procedure is used to analyze the 

scram design point of an air-breathing combined cycle 

engine. It is envisioned that the consistent 

multidisciplinary solution given by this procedure can 

aid in the engineís design, and be used for 

optimization of the design across all three disciplines. 

NASA/TMó2002-211971 9 

Appendix 

Numerical experiments and theoretical analysis 

indicate that under-relaxation may be needed for 

convergence of the fluid-thermal iteration. Some 

researchers have used under-relaxation [14,15], while 

other work has not required it [12]. The following 

theoretical analysis provides guidance for converging 

fluid and thermal solutions to a consistent solution. 

For the simple geometry of Figure A, the analysis 

derives a single equation that reflects both the solution 

Figure A: The bilayer solid contains materials A and B, with 

constant thermal conductivities kA and kB, and thickness LA

and LB. The exterior surfaces are kept at constant 

temperatures TA and TB, defining a one-dimensional 

problem. The temperature, TI, at the interface between A 

and B, is found with the same procedure as in the fluidthermal iteration. Clearly, the problem of interest is a fluidsolid and not a solid-solid interface. 

procedure and the transfer of heat fluxes and 

temperatures during a fluid-thermal iteration. The 

equation variable, ε

i

, is the difference between the 

consistent interface temperature, TI, and the calculated 

value after i iterations. 

ε

i 

 = − α ε

iñ1, where α = (kA LB /kBLA) > 0 

 

This result predicts that the iterated temperatures 

oscillate about the converged value, which we have 

observed. Further it predicts convergence is stable for 

α < 1, which corresponds to high thermal conductivity 

for the solid in the fluid-thermal iteration. Again this 

result has been observed. 

If we model the analogue of under-relaxing the 

interface temperatures calculated by ANSYS 

(parameter 0 < w ≤ 1), the equation becomes 

ε

i = [ 1 − w ( 1 + α ) ] εiñ1 

This result predicts that sufficient under-relaxation of 

the iterated interface temperatures will always yield 

convergence. In the fluid-thermal iteration, underrelaxation values, w, of 0.5 and 0.25 were necessary to 

converge the VULCAN-ANSYS and OVERFLOWANSYS iterations, respectively. 

LA LB 

Heat Flux

TA TB 

kA kB 

TI 

NASA/TMó2002-211971 11 

References 

 1. R.F. Faulkner, ìIntegrated System Test of an 

Airbreathing Rocket (ISTAR),î AIAA Paper 

2001ñ1812, AIAA (2001). 

 2. S.N.B. Murthy, E.T. Curran, (eds.), ìHigh-Speed 

Flight Propulsion Systems,î AIAA Progress in 

Astronautics and Aeronautics Series, Vol. 137, 

Washington, DC, 1991. 

 3. C. Trefny, ìAn Air-Breathing Launch Vehicle 

Concept for Single-Stage-to-Orbit,î AIAA Paper 

1999ñ2730, AIAA (1999). 

 4. P.G. Buning, D.C. Jesperson, T.H. Pulliam, G.H. 

Klopfer, W.M. Chan, J.P. Slotnick, S.E. Krist, K.J. 

Renze, ìOverflow Users Manual,î Version 1.8j, 

February 1999, NASA Langley Research Center, 

Hampton, Virginia. 

 5. F.M White, Viscous Fluid Flow, McGraw-Hill, 

1974. 

 6. ìGridgen User Manual,î Version 13, (1988) 

Pointwise, Inc., Bedford, Texas. 

 7. N.E. Suhs, R.W. Tramel, ìPEGSUS Userís 

Manual,î AEDCñTRñ91ñ8, November 1991, 

Arnold Engineering Development Center. 

 8. J.A., White, ìVULCAN User Manual,î Version 

4.3, March 2002, NASA Langley Research Center, 

Hampton, Virginia. 

 9. ìANSYS Users Manual,î Version 5.6, (1999) 

Swanson Analysis Systems, Inc., Cannonsburg, 

Pennsylvania. 

10. ìINCONEL Alloy 625,î Special Metals Corp. 

Huntington, WV, www.specialmetals.com, 

11. ìTimetal Data Sheets: TIMETAL 21S,î Timet, 

Denver, www.timet.com, 

12. A. Suresh, M.S. Liou, J. DeBonis, ìMultidisciplinary Analysis of a Supersonic Engine 

Inlet,î FEDSM2001ñ18137, ASME (2001). 

13. A. Montenay, L. Pate, J.M. Duboue, ìConjugate 

Heat Transfer Analysis of an Engine Internal 

Cavity,î 2000ñGTñ282, ASME (2000). 

14. G. Croce, H. Beaugendre, W.G. Habashi, 

ìCHT3D: FENSAP-ICE Conjugate Heat Transfer 

Computations with Droplet Impingement and 

Runback Effects,î AIAA Paper 2002ñ0386, AIAA 

(2002). 

15. J.A. White, ìA Modified Wall Matching 

Treatment to Account for Local Solid to Fluid 

Thermal Coupling,î JAANAF CS/APS/PSHS/ 

MSS Joint Meeting, April 8ñ12, 2002, Destin, 

Florida. 
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17

Multidisciplinary Analysis of a Hypersonic Engine

M.E.M. Stewart, A. Suresh, M.S. Liou, A.K. Owen, and D.G. Messitt

Propulsion; Multidisciplinary research; Hypersonics; SCRAM

Unclassified -Unlimited

Subject Category: 07 Distribution: Nonstandard

Prepared for the 11th International Conference on Space Planes and Hypersonic Systems and Technologies cosponsored by

the American Institute of Aeronautics and Astronautics and the Association Aeronautique et Astronautique de France,

Orleans, France, September 29–October 4, 2002. M.E.M. Stewart and A. Suresh, QSS Group, Inc., Cleveland, Ohio 44135;

M.S. Liou and A.K. Owen, NASA Glenn Research Center; and D.G. Messitt, Aerojet, Sacramento, California 95813.

Responsible person, M.E.M. Stewart, organization code 2900, 216–977–1163.

This paper describes implementation of a technique used to obtain a high fidelity fluid-thermal-structural solution of a

combined cycle engine at its scram design point. Single-discipline simulations are insufficient here since interactions from

other disciplines are significant. Using off-the-shelf, validated solvers for the fluid, chemistry, thermal, and structural

solutions, this approach couples together their results to obtain consistent solutions.