Ali A. Ameri

The Ohio State University, Columbus, Ohio

NASA Rotor 37 CFD Code Validation

Glenn-HT Code

NASA/CR—2010-216235

July 2010

AIAA–2009–1060
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Ali A. Ameri

The Ohio State University, Columbus, Ohio

NASA Rotor 37 CFD Code Validation

Glenn-HT Code

NASA/CR—2010-216235

July 2010

AIAA–2009–1060

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC06BA07B, Task 269

Prepared for the

47th Aerospace Sciences Meeting

sponsored by the American Institute of Aeronautics and Astronautics

Orlando, Florida, January 5–8, 2009
Acknowledgments

The author wishes to thank Drs. Rodrick Chima and Mark Celestina for their advice on this work.

Thanks are also due to Drs. James Heidmann and Ken Suder for their support of this work.

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Level of Review: This material has been technically reviewed by NASA technical management.
NASA/CR—2010-216235 1 

NASA Rotor 37 CFD Code Validation 

Glenn-HT Code

Ali A. Ameri 

The Ohio State University 

Columbus, Ohio 43210 

Abstract 

In order to advance the goals of NASA aeronautics programs, it is necessary to continuously evaluate 

and improve the computational tools used for research and design at NASA. One such code is the GlennHT code which is used at NASA Glenn Research Center (GRC) for turbomachinery computations. 

Although the code has been thoroughly validated for turbine heat transfer computations, it has not been 

utilized for compressors. In this work, Glenn-HT was used to compute the flow in a transonic compressor 

and comparisons were made to experimental data. The results presented here are in good agreement with 

this data. Most of the measures of performance are well within the measurement uncertainties and the exit 

profiles of interest agree with the experimental measurements. 

Introduction 

NASA Rotor 37 test case, designed and tested originally by Reid and Moore (Ref. 1) at NASA Lewis, 

presently Glenn Research Center, is a low aspect ratio inlet stage for an eight-stage core compressor with 

a 20:1 pressure ratio (Ref. 2). It was retested at NASA Glenn in isolation to avoid the interaction effects. 

As such the test case is ideal for code verification. The specifications, given in Table I, are the intended 

design parameters and are reproduced from Suder (Ref. 2). Much data representing performance and flow 

field variables can be found in (Ref. 2) which has been used for CFD validation. 

Given the steady nature of the flow and the isolation of the rotor blade row, the numerical simulation 

of this case should be rather straight forward. But, the results of the IGTI 1994 blind test case 

(unpublished) showed a large variation in the results and often large deviation from the experimental 

results. A new round of comparison with the data has been undertaken to assess the current state-of-theart in some of the codes used at NASA. Glenn-HT although used mostly for heat transfer computations in 

turbine blades and passages was thus assessed by computing the case and comparing with the data. 

The code Glenn-HT is a general structured multi-block code. Blocks are full face matching and can be 

arranged in an unstructured fashion. Alternatively, the computational domain can be formed from grid zones 

which are brought together at non-matching interfaces though this was not done in the present work. This 

makes the code ideal for complicated geometries and near wall solutions for which the near wall resolution 

is important. Glenn-HT uses a finite volume scheme. It uses a second order central difference scheme plus a 

fourth order artificial dissipation. A pressure switch turns on a second order dissipation that is activated by 

shocks. Thus near shock solutions are first order accurate. Glenn-HT code uses a multi-grid solver and 

utilizes a Runge-Kutta explicit smoother. Further acceleration is achieved with residual smoothing. A 

second order upwind method described by Huynh (Ref. 3) of NASA Glenn is also available. As for 

turbulence models, Glenn-HT uses Wilcox’s k-as the default model (Ref. 4). The code also uses other 

turbulence models including the v2

-f model of Durbin (Refs. 5 and 6) and a Reynolds Stress model 

developed by Wilcox (Ref. 4). Glenn-HT has found use in external and internal flows including heat transfer 

on turbine blades including tip clearance (Refs. 7 and 8), film cooling flow and heat transfer and conjugate 

heat transfer (Ref. 9). Other applications have included internal heating of airplane wings (Ref. 10). As for 

unsteady applications, solutions have been obtained for external turbine flows and heat transfer (Refs. 11, 

12, and 13) and external flows resulting from pulse detonation (unpublished). Glenn-HT code has never 

been used for compressor flows and the present application is new. 
NASA/CR—2010-216235 2 

TABLE I.—ROTOR 37 DESIGN PARAMETERS 

In an earlier incarnation, the code Glenn-HT was called TRAF3D. Arnone and Ameri using TRAF3D 

participated in the blind test case involving the present data (Ref. 14). The TRAF3D code, at the time, 

was a single block code that used a ‘C’ grid topology and employed an algebraic turbulence model. In the 

blind test case the tip clearance was modeled without gridding the gap. In the present work the tip gap 

was accounted for by meshing the tip clearance. 

Results 

In Figure 1, two views of the multi-block grid generated for the blade passage and the clearance is 

shown. The grid was generated using the GridPro commercial software. A close-up view of the tip grid is 

included that shows the refinement of the tip grid. The grid contains approximately 1.8 million cells. Only 

one grid resolution was tried but owing to the very large number of grid cells used in the simulations, no 

further refinement was done for this case. Due to the topology of the grid, the resolution of the grid on the 

tip of the blade, is as fine as the grid on the suction or the pressure side. Including the resolution of the 

grid on the casing, the tip region contains 49 grid points across the span of the tip clearance which was 

0.4 cm as reported by Suder (Ref. 2). 

Earlier studies of tip clearance effect (Refs. 15 and 16) have been made with far fewer points in the 

tip. The inlet and exit were placed so that they matched the locations recommended by the blind test case. 

The domain was subdivided into many blocks which was a result of grid generation. The blocks were 

formed into groups of approximately equal sizes in terms of grid points to take advantage of the multiprocessing capability of the code. 

For turbulence model, the Low Reynolds number k- model of Wilcox (Ref. 4) was used. Chima 

(Ref. 17) tested a newer version of the k- model (Ref. 18) and found it to perform better than the older 

version for the present case. We have not used any other models for this test case. 

The dissipation is scaled with the grid resolution in Glenn-HT. Still only small amount of fourth order 

artificial dissipation was needed to achieve stable solution and convergence. The solution at 70 percent 

span section, in terms of Mach number for the near peak efficiency condition (discussed later), is shown 

in Figure 2. The bow shock and the shock/boundary layer interaction can be seen to have been simulated 

with good resolution. On the blade suction side, shown in Figure 3, the streamlines show a small 

recirculation zone that extends from the hub to the tip. This recirculation zone results from the interaction 

of the shock with the boundary layer as seen in Figure 2. 
NASA/CR—2010-216235 3 

Figure 3.—Flow streamlines showing some important features of the flow. 

Figure 2.—Blade to blade Mach number contours, CFD (left) and experiment (right). 

 

Figure 1.—Grid for the rotor 37 computations.
NASA/CR—2010-216235 4 

Figure 4.—Streamlines showing the flow over the blade tips at the near peak efficiency (left) and 

near stall (right) conditions. 

Figure 4 shows the streamlines showing the flow over the blade tip for the near peak and near stall 

efficiency conditions. Near the leading edge the streamlines form into the tip vortex. As suggested by the 

streamlines, the tip vortex for the near peak condition is tighter and causes a smaller blockage to the flow 

than the near stall condition. Tip clearance flow has an important role to play in the performance of the 

rotor. Figure 5 shows the tip flow in the near leading edge region and in the mid-chord region. The flow 

recirculation appears to be larger in the near leading edge region and appears to diminish further down the 

blade. The resolution in the flow as can be seen is quite good. This is evidenced by the presence of core of 

the tip and edge vortices. The blade surface for the top two figures is colored according to pressure. The 

tip is seen to unload and assume the suction side pressure level very near the tip. 

Performance Measures

One and a 10 percent boundary layer thickness profiles were prescribed at the inlet to the 

computational domain. The quantities of interest were found at the inlet and the exit of the passage at 

what was referred to as stations 1 at x = 4 cm upstream and station 4 at x = 10 cm downstream as given in 

(Ref. 2). Quantities were averaged at the measuring station. 

Figure 6 shows the total pressure rise (outlet/inlet) versus the normalized mass flow through the 

passage. The map was formed by initially computing the choked mass flow for a static exit pressure ratio 

of 1.05. The static pressure ratio was subsequently raised in increments up to a value of 1.22 

corresponding to the stall condition and the exit total pressure rise and mass flow rate noted. The results 

are plotted against the experimental data provided by Suder and Celestina (Ref. 12). The experimentally 

measured boundary layer thickness is 0.82 cm which is 10 percent of the inlet passage height. The 

experimental choked mass flow was 20.93 kg/sec. The choked mass flow rate computed for a 10 percent 

inlet boundary layer thickness was 20.56 kg/sec which is 1.76 percent below the experimental value. A 

1 percent boundary layer thickness was attempted for which the computed choked mass flow was 

20.85 kg/sec or 0.38 percent below the measured value. As shown, an increase in the inlet boundary layer 

thickness causes a reduction in the choked mass flow. The resulting map, in Figure 6, for the two 

boundary layer thicknesses produce the same results for pressure ratio, temperature ratio and adiabatic 

efficiency. We will present the results for the 10 percent boundary layer thickness as in the experiment, 

but we would note here that the remaining results for the two cases are quite similar. 

Figure 6 shows the normalized total pressure, total temperature and adiabatic efficiency as a function 

of normalized mass flow rate. The agreement with the experimental data for the temperature ratio is quite 

good, but the computed pressure ratio and adiabatic efficiencies are low. 
NASA/CR—2010-216235 5 

Exit Profiles 

Two points on the map were chosen for further comparison. A near peak efficiency and a near stall 

point. The near peak efficiency was at a normalized mass flow rate of 0.98 and the near stall point was at 

a normalized mass flow rate of 0.925. We have chosen the pressure ratios of 1.12 and 1.22 as our 

corresponding points based on the mass flows and proximity to the runs we have made. 

Figure 7 shows the spanwise variation of the tangentially averaged exit (downstream station 

x = 10 cm) total pressure, for the points on the map designated as near peak efficiency and near stall. The 

near hub deficit in the total pressure, evident in the data, has been captured. The results of the 1994 blind 

test case for the near peak efficiency results are also presented for comparison. For the total temperature, 

in Figure 8, the trends are captured and the agreement is quite good. The bands in which the results would 

be considered in good agreement with the data are shown for both total temperature and for the total 

pressure. The agreement is thus quite good for both of these measures. The large over-prediction in total 

Figure 5.—Tip flow showing the recirculation zone in 

the tip clearance, top: near L.E. and middle in the 

mid chord region. The bottom figure shows the 

edge vortex along the blade tip. 

Figure 6.—Total pressure, total temperature and 

adiabatic efficiency versus normalized mass flow. 
NASA/CR—2010-216235 6 

Figure 7.—Pitchwise average of total pressure across the span (left) and the results of 1994 blind test case. 

Figure 8.—Pitchwise average of total temperature across the span (left) and the results of 1994 blind test case.
NASA/CR—2010-216235 7 

temperature near the case that was discussed by Chima (Ref. 17) has been largely eliminated. This could 

be due to the fine grid resolution used in the blade tip region. The efficiency and the exit angle are shown 

in Figures 9 and 10. The agreement with the efficiency is quite good. We have not been able to reconcile 

the contradiction presented by the agreement of adiabatic efficiency with the data shown in Figure 9 and 

lack of agreement shown in Figure 6 in the map. Finally, the exit angle distribution is also shown in 

Figure 8. There is a maximum of 3 difference between the measured and computed exit angles. 

Summary 

The code Glenn-HT was used to simulate the flow in NASA rotor 37. A fine grid with a finely 

gridded tip clearance was employed. The default turbulence model in the code namely, the k- model 

was utilized. The agreement with the data at the downstream measuring station was quite good. The 

performance map was fairly well predicted. Glenn-HT is mainly used for thermal analysis in gas turbines. 

As a code, Glenn-HT is first and foremost a flow solver. Heat transfer capabilities are added features to 

enable thermal analysis the success of which depends on the underlying flow analysis scheme and the 

associated level of accuracy. From the results of this assessment, it can be concluded that Glenn-HT may 

be confidently used for compressor flows. It is expected that even better agreement with the data can be 

attained by using more appropriate turbulence models. 

References 

1. Reid, Lonnie and Moore, Royce, D., “Performance of Single-Stage Axial-Flow Transonic 

Compressor With Rotor and Stator Aspect Ratios of 1.19 and 1.26, Respectively, and With Design 

Pressure Ratio of 1.82,” NASA Technical Paper 1338, November 1978. 

2. Suder, Kenneth, L., “Experimental Investigation of the Flow Field in a Transonic, an Axial Flow 

Compressor With Respect to the Development of Blockage and Loss,” NASA TM 107310, October, 

1996. 

Figure 9.—Pitchwise average of 

adiabatic efficiency at the exit. 

Figure 10.—Pitchwise average of 

exit angle at the exit. 
NASA/CR—2010-216235 8 

3. Huynh, H.T., “Analysis and Improvement of Upwind and Centered Schemes on Quadrilateral and 

Triangular Meshes,” AIAA–2003–3541. 

4. Wilcox, D.C., 1994, Turbulence Modeling for CFD, Third Edition, DCW Industries, Inc., 

La Canada, CA., 2004. 

5. Durbin, P.A. Pettersson Reif, B.A. Statistical Theory and Modeling for Turbulent Flow, Wiley 

Publishing, 2001. 

6. Ameri, Ali A. and Ajmani, Kumud, “Evaluation of Predicted Heat Transfer on a Transonic Blade 

Using V2

-f Models,” IGTI Meeting, Vienna, Austria, ASME 2004–GT–53572, 2004. 

7. Ameri, Ali A., Steinthorsson, E. Rigby, D., “Effect of Tip Clearance and Casing Recess on Heat 

Transfer and Stage Efficiency in Axial Turbines,” ASME Journal of Turbomachinery, 121, No. 4, 

pp. 683–693, October 1999. 

8. Ameri, Ali A. Steinthorsson, E. Rigby, D., “Effect of Squealer Tips on Rotor Heat Transfer and 

Efficiency,” ASME Journal of Turbomachinery, 120, No. 4, pp. 753–759, Oct. 1998. 

9. Heidmann, J., Rigby, D.L Ameri, A.A., “A Three Dimensional Coupled Inetrnal/External Simulation 

of a Film Cooled Turbine Vane,” ASME Journal of Turbomachinery, 122, pp. 348–359, Apr. 2000. 

10. Rigby, D. “Numerical Investigation of Hole Pattern Effect on Piccolo Tube Anti-Icing,” 

AIAA–2006–1012. 

11. Ameri, Ali A. and El-Gabry, Lamyaa, “Computation of Unsteady Heat Transfer on the Casing of a 

Turbine Blade,” ISROMAC 12, Honolulu, Hawaii, February 2008. 

12. Ameri, A.A., Rigby, D.L. Steinthorsson, E., Heidmann, J.D., Fabian, J.C., “Unsteady Turbine Blade 

and Tip Heat Transfer Due to Wake Passing,” GT2007–27550, Gas Turbine Conference, May 2007, 

Montreal, Canada. 

13. Ameri, A.A. Rigby, D.L., Steinthorsson, E., Heidmann, J.D., Fabian, J.C., “Unsteady Analysis of 

Blade and Tip Heat Transfer as Influenced by The Upstream Momentum and Thermal Wakes,” 

GT2008–51242, Gas Turbine Conference, June 2008, Berlin, Germany. 

14. Arnone, Andrea and Ameri, Ali A., “Predictions of the NASA Rotor 37 Test Case,” International Gas 

Turbine Institute Meeting presented at the ASME International Gas Turbine Institute Meeting, June, 

1994, The Hague, Netherlands. 

15. Chima, Rodrick V., “Calcualtion of Tip Clearance Effects in a Transonic Compressor Rotor,” ASME 

Paper 96-GT-114, June 1996. Also NASA TM-107216. 

16. Yamada, Kazutoyo, Furukawa, Masato, Inoue, Masahiro and Funazaki, Ken-ichi, “Numerical 

Analysis of Tip Leakage Flow Field in a Transonic Axial Compressor Rotor,” Proceedings of the 

International Gas Turbine Congress 2003 Tokyo. IGTC2003 Tokyo TS-030. 

17. Chima, R.V., “Swift Code Assessment for Two Similar Transonic Compressors,” AIAA paper 

AIAA–2009–1058 or NASA/TM—2009-215520, Jan. 2009. 

18. Wilcox, David C., “Formulation of the k- turbulence model revisited,” AIAA–2007–1408. 
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14. ABSTRACT

In order to advance the goals of NASA aeronautics programs, it is necessary to continuously evaluate and improve the computational tools 

used for research and design at NASA. One such code is the Glenn-HT code which is used at NASA Glenn Research Center (GRC) for 

turbomachinery computations. Although the code has been thoroughly validated for turbine heat transfer computations, it has not been 

utilized for compressors. In this work, Glenn-HT was used to compute the flow in a transonic compressor and comparisons were made to 

experimental data. The results presented here are in good agreement with this data. Most of the measures of performance are well within the 

measurement uncertainties and the exit profiles of interest agree with the experimental measurements.

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Computational fluid dynamics; Transonic compressors; Turbomachinery; Computational grids 

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