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Theses and Dissertations 1. Thesis and Dissertation Collection, all items

1995-09

Transonic axial compressor design case study

and preparations for testing

Reid, William D.

Monterey, California. Naval Postgraduate School

http://hdl.handle.net/10945/35185

Downloaded from NPS Archive: Calhoun

NAVAL POSTGRADUATE SCHOOL MONTEREY, CALIFORNIA

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THESIS

TRANSONIC AXIAL COMPRESSOR DESIGN CASE STUDY AND PREPARATIONS FOR

TESTING

by

William D. Reid

September, 1995

Thesis Advisor: Raymond P. Shreeve

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Master's Thesis TRANSONIC AXIAL COMPRESSOR DESIGN CASE STUDY AND PREPARATIONS FOR TESTING.

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6. AUTHOR(S) William P. Reid

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13. ABSTRACT (maximum 200 words)

Test runs of the transonic axial compressor test rig at the Naval Postgraduate School, Turbopropulsion Laboratory, were conducted in preparation for the installation of a new stage design. Modifications in the cooling air supply to the high speed bearings, and to the design of the torque measuring system were completed during subsequent overhaul.

A case study of the design of the new transonic stage was initiated. This consisted of a review of the procedure used in the design, as well as a design comparison. The comparison examined the differences between the blades designed for the new stage, which was primarily accomplished using a full, three dimensional, Computational Fluid Dynamics code, and blades designed using two-dimensional streamline curvature methods. The axi-symmetric through-flow code, used in the design case study was modified to run on workstations at the Naval Postgraduate School, Department of Aeronautics and Astronautics, providing students and faculty with a design tool for single or multiple stage axial flow compressors

14. SUBJECT TERMS

Turbomachinery AXIDES, Compressor, Transonic, CFD,

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Approved for public release; distribution is unlimited.

TRANSONIC AXIAL COMPRESSOR DESIGN CASE STUDY AND PREPARATIONS FOR TESTING

William D. Reid

Lieutenant, United States Navy

B.M.E., The George Washington University, 1988

Submitted in partial fulfillment

of the requirements for the degree of

MASTER OF SCIENCE IN AERONAUTICAL ENGINEERING

from the

NAVAL POSTGRADUATE SCHOOL September 1995

Author:

Approved by:

llliam D. Reid

Raymond P. Shreeve, Thesis Advisor

Garth V. Hobson, Second Reader

Daniel J. Collins, Chairman Department of Aeronautics and Astronautics

in

IV

ABSTRACT

Test runs of the transonic axial compressor test rig at the Naval Postgraduate School,

Turbopropulsion Laboratory, were conducted in preparation for the installation of a new

stage design. Modifications in the cooling air supply to the high speed bearings, and to the

design of the torque measuring system were completed during subsequent overhaul.

A case study of the design of the new transonic stage was initiated. This consisted of

a review of the procedure used in the design, as well as a design comparison. The

comparison examined the differences between the blades designed for the new stage, which

was primarily accomplished using a full, three-dimensional, Computational Fluid Dynamics

code, and blades designed using two-dimensional streamline curvature methods. The axi-

symmetric through-flow code, used in the design case study was modified to run on

workstations at the Naval Postgraduate School, Department of Aeronautics and Astronautics,

providing students and faculty with a design tool for single or multiple stage axial flow

compressors.

VI

TABLE OF CONTENTS

L INTRODUCTION. 1

n. NEW STAGE DESIGN 5

A-PROCEDURE 5

1. Design Intent 5

2. Preliminary Calculations 5

3. Final Selection of Preliminary Design Parameters 7

4. Blade Shape Definition 9

5. Fabrication 21

B. DESIGN COMPARISON n

1. Overview 21

2. Method Of Comparison 22

3. Results Of Comparison 14

m. PREPARATIONS FOR TESTING 21

A. OVERVIEW 21

B. PRELIMINARY TEST RUNS 21

C. OVERHAUL 22

1. Cooling Of The High Speed Bearings 23

2. Torque Measurement system 23

D. FINAL TEST RUN 30

IV. CONCLUSIONS AND RECOMMENDATIONS 31

APPENDLX A. AXD3ES CODE 35

Vll

A. OVERVIEW • • • 35

B. INPUT 36

C. CODE EXECUTION • 45

D.OUTPUT • 47

E. CONCLUDING REMARKS • • 48

APPENDIX B. AXIDES OUTPUT 59

APPENDIX C. TEST RIG 93

A. OPERATING PROCEDURE 93

B. DISASSEMBLY AND REASSEMBLY • 94

C.INSTRUMENTATION •••• 96

LIST OF REFERENCES "

INITIAL DISTRD3UTION LIST 101

vni

LIST OF SYMBOLS

w Specific Head Rise Parameter

Cp Specific Heat at Constant Pressure

To, Inlet Total Temperature

PR Pressure Ratio, Total-to-Total

y Ratio of Specific Heats

u, Rotor Wheel Speed at Tip

IX

I. INTRODUCTION

The Naval Postgraduate School (NPS) transonic axial compressor design was

completed by Dr. M.H. Vavra in July of 1968. All of the calculations and drawings

required for the complete aerodynamic and mechanical design were done by Dr. Vavra

alone, by hand. From 1968 until the mid 1980's the compressor served well as a test

vehicle for the development of innovative flow measurement instrumentation as well as a

means for graduate students to gain experience in the operation, test, and analysis of high

speed compressors [Ref. 1-5]. The initial stage design was unique but no real technology

advance was attained. This was due to two design inaccuracies, discovered after some

period of testing. In addition to an incorrect assumption of constant through-flow velocity,

an error was made in the calculations associated with the rotor blade setting angle, and

this was subsequently built into the blading [Ref. 6]. Consideration was given to twisting

the fabricated blades to correspond with the shape intended in the design. This was not

attempted since the design itself had resulted from an inaccurate through-flow prediction.

Following the completion of a study by NeuhofT[Ref 7], to measure rotor losses, and to

separately identify shock and viscous components, the compressor was not operated again

until the present study was initiated.

A transonic axial compressor stage designed specifically for the existing NPS

transonic test rig, shown in Figure 1, was completed at NASA Lewis in 1994, by Nelson

L. Sanger [Ref. 8]. This new design was sought to perform the functions originally

intended for the Vavra design, namely, to serve as a test vehicle for instructional purposes,

and to provide a tool for meaningful research. However, the timing of the project, after a

period of years in which advanced Computational Fluid Dynamics (CFD) methods for

turbomachinery had evolved, provided a unique opportunity. The design would provide a

test of a wholly CFD design approach. The experimental evaluation would provide both a

validation of the design approach, and an experimental test case for CFD analysis codes.

The manufactured blading is shown in Figure 2.

Figure 1. Transonic Compressor Test Rig.

Figure 2. Manufactured Blading; Rotor (Top), Stator (Bottom).

This report is intended to document the effort to bring the NPS transonic

compressor test rig back on line, in order to evaluate the newly designed axial stage, and

to review, for the record, the procedure used in the stage design. The report has two main

sections. The first section contains a review of the process used in the design of the new

stage. This section also contains a comparison of the final blade design, resulting from

CFD analysis, with a blade design using conventional streamline curvature and

axisymmetric through-flow calculations. The second section is an account of the process

that was conducted to return the test rig to working order. Appendix A contains a

description of, AXJDES, a 2-D blade design code, written by Jim Crouse [Ref. 9], which

was used in the preliminary design and in the production of final fabrication coordinates of

the new stage. This code has been modified for use at the NPS, Department of

Aeronautics and Astronautics, as a tool for advanced axial compressor design. Appendix

C contains information regarding the test rig disassembly and reassembly, and the

operating procedure.

II. NEW STAGE DESIGN

A. PROCEDURE

1. Design Intent

The object of the design was to produce a state-of-the-art transonic compressor

stage, representative of a typical inlet stage, or fan stage. At the time the design was being

contemplated, the use of CFD codes as analysis tools was increasing. However there was

little work done using these codes for design. Given that the stage was to be used for

research and was not intended as an actual aircraft component, the decision was made to

use CFD as the principle tool in the aerodynamic design. Additional computer codes were

used in the design including the AXIDES code, however these codes were used mainly in

supporting roles. The complete package, including the design, testing and measurement of

the stage constitutes, in the words of the designer, "the ultimate CFD validation

experiment".

2. Preliminary Calculations

Although the design was intended to be derived from state-of-the-art technology,

the decision to utilize the existing NPS Turbopropulsion Laboratory (TPL) transonic

compressor test rig imposed several constraints that were unavoidable. The decision to use

the existing test rig mechanical design was economic in nature. Retooling the rig to

produce a desired flow path was determined to be unjustified in light of the cost, and

effect on the final design. The constraints imposed on the design included a constant blade

tip diameter dictated by the existing compressor shroud, a pre-determined blade chord set

by restrictions in the flow path and the need to provide clearance for probe measurements.

A further requirement was that the stator discharge angle be axial in order to

accommodate the unique torque measuring system (see Chapter III, Section C2). An

additional constraint of significant importance was the limitation that the turbine drive unit

was capable of providing a maximum of 450 to 470 horsepower.

Within the confines of the physical constraints imposed by the existing compressor

rig, the designer conducted a preliminary analysis of desired design parameters with the

goal of producing a stage that achieved,

" as high a loading and specific weight flow as was practical, while keeping rotor tip

Mach number at a moderate level. "

Sänger'94

Using the specific head rise parameter,

T = CpTo,(PR(y-1/r)-l)/Ut2

a parametric study was performed. The specific head rise was calculated over a range of

tip speeds, and pressure ratios, using as a reference several previous designs, including the

original Vavra design and the well known NASA Rotor 67. To produce a design that was

state-of-the-art, the initial selection of design parameters was fairly optimistic. Subsequent

calculations of the horsepower and inlet ramp angle required to support this selection of

design parameters revealed that they were in fact too optimistic. The power required

exceeded that available by nearly 200 hp, and the ramp angle on the spinner hub was

calculated to be 42 Deg.

In order to reduce the power required by the stage to that constrained by the

existing drive turbine, and lower the inlet ramp angle to a more reasonable figure, several

compromises had to be made. Notable compromises were the reductions in specific weight

flow and rotor and stage pressure ratio, to alleviate the power imbalance, and a reduction

in the axial velocity ratio, to reduce the inlet-to-exit area ratio and therefore the inlet ramp

angle. With these corrections, the required input power was reduced to 457 horsepower,

and the ramp angle became 28.2 Deg. Selection of additional design parameters such as

solidity and diffusion factor was primarily based on the designer's experience with similar

machines. The results of the parametric study are shown in Figure 3.

0.4

ü

OJ

u c

00

0.3

0.2

0.1

Present Dcsig

lip Speed, m/sec (ft./sec)

365.8(1200)

396.2(1300)

426.7(1400) 445.6(1462)

NASA Rotor 67

ra Design

1.4 1.5 1.6 Pressure Ratio

1.7 1.8

Figure 3. Results of parametric study [from Ref. 8].

3. Final Selection of Preliminary Design Parameters

Prior to this stage in the design the analysis was restricted to one-dimensional

methods. To arrive at a detailed description of the flow path, and evaluate the

performance of the preliminary design, the designer used the AXIDES code. The

reasoning behind using this code was to reveal any obvious deficiencies in the preliminary

design and to produce the flow field from which the blading would be derived. The

AXIDES code provided the full 2-D radial equilibrium solution, where steady,

axisymmetric flow is assumed, and produced distributions of the flow between blade rows

at the specified design condition. At this point in the design process the flow-field

estimation option of the code was used by setting the parameter OP, of the input data set,

to APPROX (see Appendix A). Using this option, no blades are designed or specified. The

code simply estimated blade-edge locations from the stacking line, and produces velocity

diagram and performance information based on these approximations.

The information that was specified in the input data set to the AXIDES code was

RPM, mass flow rate, stage pressure ratio, inlet total pressure, inlet total temperature,

tangential velocity, the flow path of the existing compressor rig, annular station locations,

and estimated blockage. Although typically input in the form of loss-correlation tables, the

losses across the blading were implicitly defined by entering the total pressure and total

temperature distributions at the outlet of the rotor. A discussion of this option of the code

can be found in Reference 9, page 12. The total-pressure distribution was essentially

constant except for losses at the hub and tip. The total-temperature distribution was also

nearly constant, however, higher temperatures were input near the tip to account for

increased shock losses along the blades. The values used for estimated blockage and total-

pressure and total-temperature distributions at the outlet of the rotor, were based on

experience and on the values of successful stages in the same design range. Additional

information specified in the input data set were the radial distributions of blade leading and

trailing edge radii, chord length, and maximum thickness-to-chord ratio. All of these

distributions were based on the designer's experience with transonic rotors. The input

used to define the stator attributes were similar. A notable exception was the exit

tangential velocity which was zero.

Several iterations through the AXIDES code were conducted, until output

distributions of such parameters as diffusion factor and loss appeared reasonable.

Modifications to the input data file consisted mainly of variations in pressure ratio and

flow path geometry. There was initial concern about the low values of output loss

distribution across the rotor. It was decided, however, that if an error in efficiency was to

be made, to design for high efficiency and be wrong was better than designing for lower

efficiency and be correct.

At the time of this report only a representative example of the final input data set

of preliminary design parameters was available. Described as "one of the last" data sets

used during iterations through the AXIDES code, it was a very close approximation of the

actual input. This input file was run on the version of the code installed at NPS. The

output received was the designer's final preliminary design flow-field and the associated

performance parameters. Reference 10 contains this input file in its raw form, and the

associated output file. Appendix A gives a description of input and output file structure

and parameters.

Once satisfied with the 2-D flow field and performance, the preliminary design

parameters were fixed. A list of first cut, and final selection of design parameters is shown

in Table 1.

Parameter Initial Selection Final Selection Rotor Pressure Ratio 1.63 1.61 Stage Pressure Ratio 1.60 1.56 Tip Speed 1300 ft/sec 1300 ft/sec Design Weight Flow 22.21 lbm/sec 17.09 lbm/sec Specific Weight Flow 40lbm/sec-fri! SSIbm/sec-ff Specific Head Rise 0.265 0.246 Tip Inlet Relative Mach Number 1.30 1.28 Hub/Tip Radius Ratio 0.40 0.51 Rotor Inlet Ramp Angle 42.0 deg. 28.2 deg. Power Required 638 HP 457 HP

Table 1. First cut, and final selection of design parameters.

The problem now consisted of blading design to produce the loss distribution, and flow

angles, specified in the output of the AXIDES code.

4. Blade Shape Definition

The AXIDES code provided the designer with a picture of the desired flow field,

between blade rows, on the meridional plane. The next step was to fit blade shapes to the

velocity diagrams to produce this flow. From the conception of the design, it had been

decided that CFD would be used to arrive at these blade shapes. The idea was to make an

initial "guess" of blade shape, using output from the AXIDES code, run this blade through

the CFD analysis, and check the results. If the output seemed reasonable the design was

complete. If not, blade shape modification was required. The designer created initial blade

sections essentially by hand, guided by the flow angle distributions produced by the

AXTDBS code and experience with supersonic flow across blading. Iteration on blade

shape, which turned out to be required, was performed using a NASA in-house blade

element code, derived from the geometry portion of the AXIDES code.

The procedure used to screen the initial blade sections was modified over the

course of the design. Initially, a quasi-three-dimensional code, written by John Denton

[Ref. 11], was used. This code was coupled with an integral boundary layer code [Ref.

12], to check the surface boundary layer condition. The Denton code used was an Euler

code containing a simple transpiration model to estimate the boundary layer blockage.

Together these codes provided the designer with a quick tool for screening the blade

shapes.

During the process of designing the blading, the full three-dimensional version of

the Denton code (TIP3D), as reported in Denton 1986 [Ref. 13], became available. This

code was simply an upgrade in the original Euler code, containing a simple approximation

for viscous effects and an empirical equation defining the distribution of shear stress from

the wall. Consideration was given to using a code with a more complex and possibly more

accurate turbulence-modeling scheme, however, given the iterative nature of the design

process, the increase in accuracy was deemed insignificant when compared with the

increase in computational execution time.

Using the quasi-3D results as a "base design", the iteration on blade shape was

continued in the same manner as previously described, using the updated version of the

Denton code. The criteria used for acceptance of the blade included the requirements to

minimize shock strength in the tip region and to prevent or delay boundary layer

separation. The shock strength was reduced by minimizing the supersonic acceleration of

the flow approaching the shock. This was accomplished by making the suction surface at

the leading-edge portion of the blade wedge shaped, and curving it after the shock to

10

provide the required turning. To prevent separation, the attempt was made to control the

diffusion towards the trailing edge of the blade. Final-design blade sections, used to

generate the now-rotating shape for fabrication, labeled "Sänger Design", are shown in

Figure 7, located in section B3 of this Chapter. Because these sections were used for

fabrication, they represented elements of the blade on cylindrical planes, parallel to the axis

of rotation of the rotor.

5. Fabrication

After all criteria had been met, to within certain acceptable tolerances, the

aerodynamic portion of the design was complete. What remained was the mechanical

analysis and fabrication of the blading. Historically, the Engineering Design Department at

NASA-Lewis had used the same geometry specifications as prescribed in the AXIDES

code, to perform their mechanical analysis. Therefore, the final blade parameters were

input back into the streamline curvature code using the output option (OP), COORD (see

Appendix A). This option created an output file containing blade coordinates for

fabrication. A copy of the input file used to produce the final blade fabrication coordinates

and the output file containing these coordinates can be found in Reference 10.

B. DESIGN COMPARISON

1. Overview

Streamline curvature methods have been used for the design of many successful

machines. Historically, the flow-field between blade rows is calculated using these

methods, a blade shape is arrived at, and the flow within the blade row, where any

problems are likely to occur, is checked by a more detailed analysis code. The AXIDES

code was written with the objective of integrating the design and analysis portions of the

overall design process in a more efficient manner. It produces through-flow output that

can be directly input into the blade-to-blade analysis codes T-SONIC [Ref. 14], and

MERIDL [Ref. 15]. The code also provides input features which allow for corrections in

11

the blade shape, determined as necessary from the output of these analysis codes, to be

easily made. The result is a composite through-flow code, capable of both aerodynamic

and blade design, that can be interfaced with two codes that analyze the flow within the

blade rows. Currently, the version of the AXIDES code at NPS, is not capable of

interfacing with MERIDL. An update to the code should include this feature.

For the reasons mentioned previously, CFD techniques were used throughout for

the design and analysis of the new Sänger stage. The AXIDES code was not used as a

blade design tool. It was simply used to describe the flow-field which satisfied the

requirements dictated by the preliminary design parameters. No blade shapes were input,

nor specified by the code, and only velocity diagram and performance information was

output. The initial "guess" of the blade shapes was generated by hand. Although iteration

on blade definition was performed using a geometry routine extracted from the AXIDES

code, the blade sections were stacked based on information produced by the CFD analysis,

and not within the aerodynamic iterations of the streamline curvature method. This section

contains a comparison between the final Sänger design and a design produced strictly with

the use of the AXIDES code; that is, the blading specification and aerodynamic

performance output by AXIDES, when used in a design capacity. The motivation for this

comparison was two-fold. First, the Crouse 2-D code was to be used for instruction in

compressor design at NPS. A comparison with the geometry arrived at using advanced

analysis methods, was a useful exercise of the code. Second, no similar comparison was

found between a design resulting from relatively simple, and computationally-efficient

streamline curvature methods, and the geometry obtained using more detailed flow

analysis routines, requiring significant investments in computational time.

2. Method Of Comparison

The performance of the Sänger design was taken from Sänger, 1994 [Ref. 8]. The

blade definition parameters and performance distributions, as described previously, were

set using the Denton TIP3D code. The AXIDES code was used strictly to provide the

12

initial flow-field estimate, and produced coordinates of the final blade design in a form that

could be conveniently used in the mechanical analysis and the subsequent fabrication

process. During the design review, when the picture of the procedure used in the design

became clear, there was initial skepticism as to how the final blade design parameters

could be input into the AXIDES code, simply to provide the desired form of the blade

geometry in the output. The concern was that the program would modify the blade shape

as it attempted to converge on a 2-D radial-equilibrium solution. A solution which would

obviously be, at the least, slightly different than the results of full 3-D CFD calculations.

This concern was reconciled by checking the blade geometry parameters in the final output

of the AXIDES code, that was sent to the NASA-Lewis Engineering Design Department,

against the results of the CFD design, as reported in Sänger '94 [Ref. 8]. Specifically, such

blade parameters as, radial distribution of maximum thickness, and maximum thickness

location, extracted from the AXIDES output, were reviewed for discrepancies; none were

found. The blade shape produced by the AXIDES code was identical to the blade

specified in the input. This appeared to have been accomplished by completely specifying

all blade parameters in the input file. The aerodynamic output was of no interest to the

designer. Aerodynamic input data were used simply to fill in the spaces of the input data

set and did not represent the performance generated when the code was run. Again, the

AXIDES code was used simply to reproduce the geometry of a blade designed using CFD

in the form required by the NASA mechanical design system.

The design produced using the AXIDES code, and henceforth referred to as the

"conventional design", was to represent what the Sänger design would have looked like if

it had been sent for mechanical analysis directly after the initial design parameters were

fixed. The blades were designed by taking the final form of these initial parameters, which

Sänger used to build his "guess" of the blade shape, and inputting them back into the

AXIDES code. In this case, however, the input file was modified to produce blade shapes

for fabrication instead of merely calling for flow-field description, as had been done in the

Sänger design. This was accomplished by changing the OP parameter to COORD and

13

using defaults for incidence angle and deviation angle by choosing 2-D for input

parameters AA, and BB. The 2-D option resulted in the use of calculations specified in

NASA SP-36 [Ref. 16], for these angles. Additional modifications to the input file

included setting option CC to OPTIMUM, which assigned a value to the turning rate ratio

of the blade element segments using an empirical function of inlet relative Mach number;

setting option DD to shock, which placed the transition point at the location of the shock;

and setting option EE to TRAN, which located the maximum thickness point at the

transition point. A copy of the input file, and output file are located in Reference 10. A

detailed description of the AXDDES code available input options, can be found in

Appendix A Section 2, and Reference 9.

3. Results Of Comparison

As previously noted, the intent of the Sänger design was to produce blading with

increased loading at lower tip speeds. A plot of rotor diffusion factor, indicative of loading

on the blade, is shown in Figure 4. It is clear from this plot that the loading on the Sänger

rotor is larger over 90% of the blade than that of the "conventional" design. Using the

detailed picture of the flow-field provided by the CFD code, the designer was able check

for flow separation along the suction surface of the blade as he increased the amount of

turning imparted to the flow. Figure 5 shows the amount of energy addition provided by

the rotor in the form of total temperature ratio. Figure 6 illustrates the increase in total

pressure ratio across the rotor. The plots indicate an increase in total pressure ratio and

energy addition produced by careful analysis of the flow within the blade row. The blade

sections for the conventional design and the Sänger design are shown in Figure 7. This

figure displays the blade sections at 90%, 50%, and 20% span, from the hub. Looking

closely at these sections, it can be seen that the curvature on the suction surface of the

Sänger rotor, near the leading edge, is very small. This was done to minimize the

supersonic acceleration of the flow as it approached the shock, and therefore reduce its

strength. The AXIDES code used in the design of the conventional blade also attempted

14

to account for the effects of a large inlet relative Mach number. This was accomplished by

activating the blade definition option OPTIMUM, located in the input data set. It can be

seen, however, that the curvature in the leading-edge portion of the conventional blade,

although modified, is still larger than on the Sänger blade. The turning of the flow was

accomplished further aft on the blade, as can be seen in the distribution of blade maximum

thickness location shown in Figure 8. The losses across the rotor were kept nearly the

same as the conventional blade. A plot of the distribution of loss coefficient is shown in

Figui re 9.

it« \ \ \ \

\ X i« - \ \,

7» - \ >v

s

■

* 4. \ \

\ \ )• - \ \

2* - \ \

1» -

\ )

1 - ^S^ /

• .) 1 ' i i J

* ••* ••«» »44 *.4t 1.4 t

D-Fact«r —^—.

Figure 4. Rotor diffusion factor.

15

s

I* 5

— S ii|tr D t •!( » - C ■■>>• Hull

1 .2 2

Figure 5. Total temperature ratio across the rotor.

i s !

* *

11 <

»i

'Sn|it D • ll| ■ C • ■ Y c ■ II. . ■ I

Figure 6. Total pressure ratio across the rotor.

16

S*n|«r DttlfN

Stag" Drtif

Scalt (In.)

Caavcaiiwiii Blttft

• o.s

Figure 7. Tip, mean-line and hub blade sections (from top).

17

100 -

90

80

70

60

50

40

30

20

10

0

/

0.35 0.45 0.55 0.65 0.75

Max. Th. Pt. Loc./Chord

Sänger Design i

I Conventional

0.85

Figure 8. Rotor maximum thickness point location.

100 T

90

80 /^

3

£ o 'tit

70

60

50 if e m a. m 40 \ \ *

30 \ \

20 -

10

0 \ >

D 0.02 0.04 0.06 0.08 0.1 0. 12

Loss Coefficient

'. Sänger Design; — - Conventional

Figure 9. Rotor loss coefficient distribution.

The preceding comparisons show that the Sanger blading inputs a larger amount of

work than blading designed using 2-D methods alone, while maintaining very similar

through-flow and loss profiles. Assuming that the CFD code used in the design predicts

the 3-D flow field accurately, the design represents a step forward toward optimizing the

design of highly loaded, efficient, transonic blading.

19

20

III. PREPARATIONS FOR TESTING

A. OVERVIEW

The transonic stage designed at NASA-Lewis, was designed to be installed in the

NPS/TPL transonic compressor test rig, which is shown schematically in Figure 1. Since

the rig had not been operated since 1986, an evaluation of its condition, and overhaul of

some its components was necessary. This section gives an account of the preliminary

evaluation and subsequent overhaul, to date, required to prepare the machine to test the

new design.

B. PRELIMINARY TEST RUNS

Preliminary test runs of the transonic compressor test rig were conducted in early

1995. The tests were run with the existing machine as it was left, with the initial rotor,

designed Dr. Vavra, still installed. This testing was conducted to become familiar with the

operation of the test rig and to determine the extent of overhaul necessary to return it to

routine working order. The operating procedure is given in Appendix C.

A total of four tests were run over a period of two months. The first test was

considered a basic operational check to make sure the machine would turn over with no

major component failure. The rig was accelerated up to 5,000 RPM, un-throttled, with the

inlet piping removed to prevent any deposits in the piping from being pulled into the

blading. During this test the following mechanical aspects of the machine were monitored,

1. Pressure ratio across the laboratory 12 stage axial compressor supplying

the drive turbine.

2. Balance air used to provide relief of the axial load on the high speed

compressor bearings.

3. Test rig rotational speed.

4. Drive turbine and compressor bearing temperatures.

air to

21

All of these indications were monitored at the test control panel. A photograph of the

control panel is given in Appendix C.

The results of the operational check were encouraging. Although, the test rig had

not been operated for many years, there was no appearance of malfunction or unusual

operation. The tests that followed were conducted in the same manner, however the RPM

was increased with each run, to a final value of 20,000 RPM. During the testing, the

bearing temperatures were monitored and compared with values recorded from test runs

conducted years earlier. These values remained nearly identical to those previously

recorded until approximately 15,000 RPM (50% design speed). Above 15,000 RPM the

bearing temperatures increased at a greater rate than previously documented. By adjusting

the oil-mist cooling air and oil supply settings, the temperatures of the bearings supporting

the turbine-drive shaft and the downstream end of the compressor shaft were controlled to

within 2 degrees of earlier recorded values. This was accomplished by reducing the oil

mist drop rate from approximately 80 drops per minute to 12 drops per minute, and

increasing air supply pressure from 20 psi to 35 psi. The bearings supporting the upstream

end of the compressor drive shaft, however, failed to respond to the changes in the

cooling system. They remained approximately 25 degrees high and continued to increase

at a rate of approximately 1 degree per hour.

The decision was made to tear down the compressor in an effort to resolve the

problem with cooling of the compressor inlet bearings, and to check the bearings for any

damage that might have occurred due to the high temperatures. Additionally, a decision

was made to leave the turbine-drive unit intact since no problems in its operation had been

revealed.

C. OVERHAUL

The transonic compressor test rig was a unique machine that had undergone

several modifications over the years. Although some modifications were documented, the

22

overhaul that was conducted for the present project revealed that some of the changes

made to had not been recorded. Additionally, the procedure involved in disassembling and

reassembling the test rig was undocumented. The overhaul presented a significant

challenge because of the mechanical complexity inherent in the arrangement of a high-

speed air turbine driving a high speed research compressor. The goal was first to

determine the method by which the rig was to be disassembled and then to carefully

disassemble and reassemble without causing damage to any of the mostly non-standard

components. During the course of the tear down and rebuild effort, video tape was taken

to document the procedure.

1. Cooling Of The High Speed Bearings

The driving consideration behind the overhaul was the problem encountered with

bearing cooling. Disassembly of the compressor hub revealed that, of the two cooling lines

intended to provide oil-mist air to the compressor inlet bearings, one was completely

disconnected. A second line which had not been indicated on the original design drawings,

was bent so that no air could flow through it. Figure 10 shows a drawing of the

compressor hub assembly. The bearings, notably the ones to the right, were not receiving

the oil-mist air supply needed to cool them. To correct the problem the cooling lines were

fitted with new tubing. As an indication of previous problems with bearing temperatures, a

one-eighth inch copper water cooling line was found encircling the inner compressor hub.

This water cooling device was not shown on the original drawings, nor recorded in any

available test log.

2. Torque Measurement system

The torque acting on the test rig rotor drive shaft was measured by the deflection

of a cantilever beam. The beam extended between the outer portion of the compressor

hub, that rotated on bearings, and the inner hub annulus which was stationary (see Figure

10). The rotation of the outer hub was a result of the swirling air-flow exiting the rotor

23

0~yyy> ■//:y/r,//kAs//4</?'/'/.- '.:A--:-y.-y//^.\ ^^^ty/*.-/.y/'''Zw'/s/- ■■■■■y■v.-///////,A

^r^%

Figure 10. Compressor hub assembly.

24

and impinging on the stator. The stator removed most of the swirl component of the

velocity, and a honeycomb flow straightener, removed the residual. The stator and flow

straightener were both mounted on the outer hub which rotated on the outer hub bearings.

The cantilever beam (torque balance), whose fixed end was attached to the outer hub and

free end was constrained by a channel fixed to the inner stationary hub annulus, deflected

as it tried to prevent relative rotation. Four strain gauges, mounted on the cantilever beam,

provided the signal that was a measure of the torque imparted to the air-flow by the rotor.

A static calibration of the torque balance was conducted in preparation for the

tests to follow overhaul. A significant drift in the strain gauge output was found. A plot of

the calibration is shown in Figure 11. The calibration was performed using weights hung

from a 20.01 inch moment arm which was mounted on the outer hub. The load was

increased from zero to 35 pounds in five pound increments. A zero of-1.0 milli-volts was

set on the output voltage display. This was done because the beam appeared to have some

pre-load on it and the bridge output could not be set to zero. The plot clearly indicates

repeatability, however, the measurements, conducted for two static load and unload

cycles, were taken after the output settled down after a drift of approximately two milli-

volts per second. This required waiting an average of 15 minutes from the time the load

was changed until the time the readings were taken. The time dependent nature of the

static load, as well as other difficulties with the torque measurement, were documented for

an early compressor test program using the same rig, by R.P. Shreeve [Ref. 17].

After careful consideration of the problem, and the ramifications of possible

changes to the system, modification of the torque balance assembly was made as shown in

Figure 12. The first change was an attempt to insure the fixed end of the cantilever beam

was firmly "encased" in the outer hub. Four set screws, mounted in the outer ring were

installed to apply pressure to the top of the beam and secure it in place. They replaced the

single (center) set screw in the original design. Additionally, the channel fixed to the inner

hub annulus was modified. The side of the channel resisting the load was drilled and fitted

25

with a three-eighths inch ball bearing. This was done to insure a point load would be

transmitted at the free end of the cantilever beam.

Strain Gauge Output (mv)

1 O 0

r~ r r- r~ 3

a:

3 a

o a.

o a> Q.

~" ' a 3

a o tu <T>

ai

O CD

3 o 3

3 CO

a> </>

Figure 11. Torque balance calibration (first).

26

Set Screws Rotation

Inner Hub Annulus (Stationary)

Outer Ring

Figure 12. Torque balance modifications.

27

A second calibration of the torque balance was conducted in the same manner as

the first with the exception that the zero-load output was set at 0.1 milli-volts. A plot of

the calibration is shown in Figure 13. The results indicated that the drift in the output

reading had disappeared. However, as can be seen clearly in Figure 13, there was a distinct

lack of repeatability over the four load and unload cycles. The output seemed to settle out

after the second cycle, however, the overall uncertainty for the four data sets was in the

vicinity of 10%. Since the uncertainty in the torque measurement gives an equal

uncertainty in the measurement of efficiency the lack of repeatability is unexceptable.

However, since the intent of the overhaul effort was to return the rig to stable mechanical

operation, no further modifications were attempted. Recommendations for additional

modifications to the torque assembly include,

1. Installing additional set screws to attach the outer ring firmly to the outer hub,

to prevent movement of the parts constraining the fixed end of the cantilever

beam.

2. Fitting the existing beam with new strain gauges.

3. Replacing the existing beam and strain-gauges with a new design having a

larger, rectangular fixed end.

28

c. n> o> o u

0) c a re CD ■n n

o o Ü V _i _i

£ o r r T3 T3 re CO re re £ g

o o

o o 00

O + X

co o o d ii

o o 03

° s

(Aiu) »ndjno »6neo uiejjs

Figure 13. Torque balance calibration (second).

29

D. FINAL TEST RUN

A final test run of the transonic compressor was conducted up to a speed of

20,000 RPM. This testing was performed un-throttled, with the inlet piping removed. The

test rig setup was similar to the preliminary test runs except for the addition of a

honeycomb flow straightener and the original stator that were installed during overhaul.

The duration of the test was two hours and twenty minutes. The objective of the test run

was to check the mechanical operation of the rig, and in particular, the behavior of the

bearing temperatures, following overhaul.

The test rig operated smoothly during the entire run, at no time indicating

problems with balancing or vibration. At a speed of 15,000 RPM, the temperature of the

bearings supporting the inlet end of the compressor drive shaft were stable, but reading

approximately ten degrees high, relative to temperatures recorded during earlier test runs

at the same RPM. As the rotational speed was increased from 15,000 RPM, to 20,000

RPM, the temperature of these bearings increased significantly. The temperature was

recorded to be 143 Deg. F, and climbing at a rate of approximately five degrees per

minute, when the RPM indicated 18,000. All other bearing temperatures were stable and

within limits. The temperature limitations of these bearings require operating levels of less

than 160 Deg. F. The RPM was reduced to 10,000 RPM and after temperatures stabilized

the machine was shutdown.

30

IV. CONCLUSIONS AND RECOMMENDATIONS

A case study was initiated of the design of a new transonic compressor stage

which is to be tested at the Turbopropulsion Laboratory at the Naval Postgraduate

School. The stage was designed by Nelson Sänger of the NASA Lewis Research Center,

using 3-D, CFD methods. The program of testing, scheduled to follow installation of the

new stage, will provide data with which to validate those methods used in the design, and

serve as a test case for still-emerging analysis codes.

The process followed by the designer was reviewed and a version of the code

(AXIDES), which was used early in the design was installed and made operational at the

School. The code was used successfully both to examine the design process for the new

stage, and later to carry out the design of a multi-stage core compressor. In preparation

for testing, the test rig was operated, partially overhauled, and operated again to 75%

design speed. The following were concluded:

Stage Design

• The initial flow-field used by the designer was reproduced successfully using AXIDES,

and the (conventional) blading associated with this flow-field was generated.

• More highly loaded blading was produced using CFD advanced design methods than

could have resulted from the use of the AXIDES code alone.

• The differences in blade shape that enabled the higher loading were, visually, relatively

subtle.

• The new stage was manufactured successfully and is ready to test after installation in the rig.

31

AXIDES Code

• The AXIDES code was made operational on NPS, AA Department workstations, and

on the TPL PC.

• The code was run for several design cases, the input and output of which serve as

examples of the varied uses of the code.

Test Rig

• Preliminary test runs were conducted up to 20,000 RPM (design speed is 30,000

RPM).

• The machine was overhauled to correct problems with cooling of the compressor inlet

bearings. The bearings were in good condition and were re-installed. The cooling air

supply system was checked and returned to working order.

• The torque measuring system was modified; however, a lack of repeatability remains.

• Following the reassembly, a test run was completed up to 20,000 RPM, and indicated

continuing problems with the bearings at the inlet end of the compressor.

• The pressure at the rotor exit is greater when the machine is run without the inlet

piping, than when the piping, screens and throttle are installed. Since the rotor exit is

where the cooling air is exhausted after passing through the bearings, the higher

pressure may be preventing the cooling air from circulating as intended.

32

The following recommendations are made:

• The "conventional" blading design should be put through CFD analysis, using a more

recent 3-D code, and compared with the results of the Denton, TIP3D code.

• The tools for complete axial compressor design, including preliminary 1-D, 2-D, and

3-D CFD, codes, all now available at NPS, should be integrated.

• The AXIDES code should be upgraded to include a graphics package, and reliable

interfacing with the off-design code AXIOFF.

• The transonic test-rig inlet pipe and housing should be installed, and the test rig

operated, to determine if the reduced pressures at the rotor exit effect the compressor

inlet bearing temperatures.

• The new stage testing program should move ahead without delay to prevent the recent

experience from being lost.

33

34

APPENDIX A. AXIDESCODE

A. OVERVIEW

The AXIDES code written by Jim Crouse is a composite aerodynamic and blade

design code. It uses the streamline curvature method to arrive at a full, 2-D, radial

equilibrium solution of the flow between blade rows in the meridional plane, assuming

steady, axisymmetric flow. The code can be used for either aerodynamic or blade design,

or both. In the case where both design capabilities are required, the code performs four

iterations of the aerodynamic solution, and then stacks the blades. This process is

continued until the solution converges. The final blade shape is calculated after the

aerodynamic solution is complete. The code can be used for either single or multistage

axial compressors. A complete description of the AXIDES code can be found in

Reference 9.

A version of the AXIDES code was purchased by the Naval Postgraduate School.

The code is originally PC based and has been modified to run on the Department of

Aeronautics and Astronautics workstations. The PC version was installed and operated on

the Turbopropulsion laboratory's 486 machine. After overcoming initial difficulties with

input and output assignments and initialization of variables, the code was used for the

following,

1. To reproduce the initial flow-field solution used in the design of the new, NPS

transonic compressor stage.

2. To reproduce the geometry of the final blade design, of the same stage, used in

the mechanical analysis and fabrication of the hardware.

3. To design a stage using the strictly 2-D calculations of the code, for the

comparison with the new stage, arrived at using 3-D CFD methods.

4. To design a multi-stage compressor for the AA Department Engine Design

Course.

35

B. INPUT

Creating the input file to the AXIDES code is crucial to its successful use as a

compressor design tool. Along with the need for a basic understanding of internal and

turbo-machinery flows, a general understanding of the input file structure is required to

produce realistic output. This section provides the tools necessary to build a working input

file.

The input to the code is in the form of a formatted data field. Each piece of the

data has a specific location in the data field and a specific length. If the input data is not in

the correct location, the formatted READ statements in the code will not read the correct

information and the program will run based on invalid input, or may not run at all.

Note

An important note here is that a "dummy" editor, such as Notepad ,on the PC, or

jot, on the workstations should be used to create or modify the input file. Other

editors including DOS and word processing editors have been found to change the

formatted structure of the file, however slightly, enough to move the input data out

of position and prevent the correct input from being read.

Figure 14, located on page 49, shows an example working input data file. This file

was provided with the code in an effort to get the program up and running on the AA

Department machines. To arrive at a successful input file it is recommended that an

example file be run first, to become familiar with the operation of the code, and then

modifications to that file be made to create an original design. Five input files were used to

produce this report. Four of these, namely,

36

• Examp.dat : The example file provided when the code was purchased,

• Approx.dat: The file used to set final preliminary design parameters, and

define the desired flow-field for the new Sänger stage,

• Conv.dat : The file used to produce the "conventional" stage for the

design comparison,

• Multax.dat : The file used in the design of a four stage compressor,

are provided for use as a starting point to build an input data set. The input file Examp.dat

is shown in Figure 14. A copy of the output associated with Examp.dat (Examp.out), is

located in Appendix B. The last three files listed above can be found in Reference 10. The

fifth input file, also located in Reference 10, titled Sanger.dat, was used simply to produce

fabrication coordinates and the aerodynamic input is not representative of the blades that

the code designed; See Chapter II, Section B2, paragraph 1, for a discussion of this.

The input file to the code is broken up into two main sections. The first section

contains general information which includes the number of stages, the flow path or casing

and hub geometric coordinates, the RPM, and the desired overall pressure ratio. The

second section consists of annular calculation station and blade row data. This data

includes axial locations of annular stations and blade rows, boundary layer blockage

estimates, the number of blades on the rotor and stator, solidity, and blade definition

parameters such as leading and trailing edge radii.

Figures 15-19, located on pages 50-54, provide another look at the example input

data file mentioned previously (Examp.dat). This illustration is included as an aid in

identifying the input file parameters used for this particular design. It must be emphasized

that this figure is not a working file but is included only for parameter clarification. The

corresponding working input file is shown in Figure 14. Also included is an illustration of

the general form of the input (Figures 20, 21, and 22, pages 55, 56, and 57), taken from

notes provided by the author of the code, with all possible input parameters and associated

formatting. For a complete description of the code and its input see Reference 9.

37

The following is a description of the input parameters used in the example input

file Examp.dat. It addresses each parameter in the file in the order shown in Figures 13-17.

The information provided below was extracted entirely from Reference 9, and is included

merely for the readers convenience.

Indices

• I

• IROTOR

• J

• K

Calculation station index

Rotor index.

Streamline index. Numbered from one at the tip.

Loss set index.

General Input

• Column

• TITLE

• ITG

Not an input parameter. The first three lines simply show the

relationship of the data below to the 80 column FORTRAN data

field.

The title of the design.

Output blade fabrication coordinate specification. The output of the

blade shapes produced by the code is in the form of suction and

pressure surface coordinates on blade sections. These coordinates

are located at uniform and round number increments of the

chord-wise distance from the leading-edge. Included in the table of

coordinates are the tangency points between the blade surfaces and

the leading and trailing-edge ellipses. The locations of the point of

tangency for the suction and pressure surfaces will be different. If

the parameter ITG is set to 0, the program will insert an

artificially high value of 99.9999, for the height of the surface

38

ITG (cont'd)

IBR

NSTRM

NROW

NA

NLOSS

NTIP

NHUB

ROT

FLOW

PRATIO

MOLWT

opposite the surface that is tangent at that chord-wise distance. This

clearly identifies the junction point between that surfaces and the

end ellipse. The leading and trailing-edge information is provided

below the table of coordinates. If ITG is set to 1, the program

simply outputs the actual height of the surface opposite the surface

that is tangent at that point. The tangency points must then be

determined by noting the change in the surface heights near the

leading and trailing-edges.

Set equal to 0 to get error messages printed when running the

code. Set equal to 1 to suppress the error message output. If set

to a value greater than 1, the program will provide no error

messages during the run, and stop after the performance summary

is printed, printing neither radial distribution information, nor

fabrication coordinates, even if specified in the input file.

Number of streamlines (11 max.).

Number of blade rows (20 max.).

Number of annular stations. Fifty maximum, including leading and

trailing edges of each blade. Must be at least four prior to first blade

row and three after the last.

Number of loss sets input in the form of tables of correlation data.

Number of outer case wall geometric coordinates.

Number of hub geometric coordinates.

Compressor RPM.

Mass flow rate (lb./sec).

Desired overall pressure ratio.

Molecular weight of air.

39

• SCALEF

CP(K)

LUNITI

LUNITO

FLO(K)

TO(l,J)

PO(l,J)

VTH(1,J)

XTIP(I)

RTIP(I)

XHUB(I)

RHUB(I)

DLOSS(K,J,l)

DFTAB(K,J,1)

Scaling factor. If set to other than 1.0, the program will scale all

linear dimensions and the weight-flow to a different size

compressor, by the factor indicated.

Constants for specific heat polynomial.

Units of input. Set equal to 1 if SI and 2 if English.

Desired units of output. Set to 1 for SI and 2 for English.

Cumulative flow fraction at streamlines, from the tip.

Inlet total temperature at streamlines from tip.

Inlet total pressure at streamlines from tip.

Inlet tangential velocity at streamlines from tip.

Axial coordinate of casing wall points.

Radial coordinates of casing wall points.

Axial coordinates of hub wall points.

Radial coordinates of hub wall points.

Array of loss parameter correlations (five sets max.).

Array of diffusion factor correlations (five sets max.).

Calculation Station And Blade Row Data

• AA Type of axial station :

ANNULAR

ROTOR

STATOR

,or, depending on location in data set,

Incidence angle input

: 2-D NASA SP-36 correlations performed internal to

code.

: 3-D NASA SP-36 correlations.

40

• AA (cont'd) : SUCTION Zero inc. to suction surface.

: TABLE Input in tabular form as INC(IROW,J), at the end of

associated blade rows input

data set.

• ZTIP(I) Tip axial coordinate of annular station or blade stacking line.

• ZHUB(I) Hub axial coordinate of annular station or blade stacking line.

• BT(I) Tip boundary layer blockage for annular station or blade LE or TE.

Can be input as a fraction of annular area, or as the displacement

from the wall in inches if preceded by a negative sign.

• BH(I) Hub boundary layer blockage (same as BT except applied at hub).

• BLEED(I) Fraction of weight flow bled off at annular station or blade row.

DLIM(IROW) Diffusion factor limit. Applies to tip for rotors and hub for stators.

If limit exceeded the program will reduce the energy addition across

that particular blade row and try to make it up in others that

have not exceeded their limit. If all blade rows are at there

limit the overall pressure ratio will be reduced.

• ALIM(IROW) Minimum relative flow angle limit leaving rotor hub, or, maximum

allowable Mach number entering the stator hub. Adjustments made

to satisfy limitations are the same as for DLIM except applied to

flow angle and Mach number.

• CRENGY Cumulative fraction of energy input across rotor to that input

across stage. If greater than 2, interpreted as rotor tip exit total

temperature in degrees Rankine, and is converted to energy

addition internal to program using the total temperature profile

input in the form of parameters PRA-PRE.

• BMATL(IROW) Material density of rotor (lb./in.3).

• NXCUT(IROW) Number of blade sections for which fabrication coordinates will be

produced. If zero, the program will select this number based on

41

NXCUT (IROW) aspect ratio. If negative, represents the number of sections desired,

(cont'd) and program reads specific locations in a table input as parameter

XCUT, at end of entire data file.

ILOSS Loss set used for blade row. If less than or equal to 0 total

pressure at the blade row exit is input instead of losses being

computed internal to the program. See Reference 9 for a discussion

of parameters PTT and PTC.

OP Input option controlling amount and type of output information.

: APPROX Only velocity diagram output based on estimated

blade edge locations.

: VEL. DIA. Only velocity diagram output based on blade edge

locations that are input as ZTEMP, and RTEMP, at

the end of the associated blade rows input data set.

: DESIGN Output consists of velocity diagram information

only based on stacked blade edge locations

computed internal to code.

: COORD Output includes velocity diagram and blade

section fabrication coordinate information

based on stacked blade edge locations.

AB Part of incidence angle TABLE input option. If SS, the code will

reference incidence angle input to the suction surface at the leading

edge. Otherwise it will be referenced to the leading edge center-

line.

BB Deviation angle input.

: 2-D NASA SP-36 correlations.

: 3-D NASA SP-36 correlations.

: TABLE Input in tabular form as parameter DEV(IROW,J),

at the end of associated blade rows input data set.

42

BB (cont'd)

• CC

DD

• EE

OPTIMUM

: TABLE

: CARTER Deviation calculated internal to the code

using Carter's rule.

Blade element geometry input.

: CIRCULAR Code Designs circular arc blade sections.

Curvature at the leading edge set by an

empirical function of inlet relative Mach

number. Below M'l of 0.8 the shape will be

circular arc. As the relative Mach Number

increases, the ratio of turning rates of the

front and rear blade segments is reduced to

avoid large shock losses.

Ratio of front to rear segment turning rates.

Input in tabular form as PHI(IROW), at the

end of associated blade rows input data set.

Input location of transition point.

: CIRCULAR Transition point between front and rear

segments put at mid-chord.

Transition point located at suction surface

shock attachment point.

Input as TRANS(IROW,J), in tabular form,

at the end of the associated blade rows input

data set.

Maximum thickness location input.

: TRAN Locates the max. thickness point at the

transition point.

: TABLE Maximum thickness location input as

ZMAX(IROW,J), in tabular form, at the end

of the associated blade rows input data set.

SHOCK

: TABLE

43

• EB

CHOKE(IROW)

• BLADES(IROW)

• SOLID(IROW)

• TILT(IROW)

• PRA-PRE(IROW)

TALE-TDLE(IROW)

• TATE-TDTE(IROW)

TAMAX-TDMAX(IROW)

Used with TABLE option of max. thickness

location. If LE, the maximum thickness location is

input as a fraction of blade chord from the leading

edge. Otherwise it is input as a fraction of the chord

from the transition point.

Choke margin input. If greater than 0 the program

will adjust incidence angle in an attempt to provide

this margin. If 0 the code makes no adjustment.

Number of rotor or stator blades.

Input solidity.

Stacking axis tilt angle.

Coefficients of the input pressure or temperature

profile polynomial at the exit of a rotor, or

tangential velocity at the exit of a stator. The

pressure profile is input here if stage energy addition

is input using CRENGY. If the rotor tip exit total

temperature profile is input as the parameter

CRENGY, the total temperature profile polynomial

coefficients are input here and the pressure profile is

input using PTT and PTC.

Polynomial coefficients describing the distribution of

the ratio of blade element leading-edge radius to

chord.

Same as TALE-TDLE except applied to the trailing-

edge.

Polynomial coefficients describing the distribution of

the ratio of blade element maximum thickness to

chord.

44

CHORDA-CHORDC(IROW)

IDEF(IROW)

• INC(IROW,J)

• DEV(IROW,J)

• PHI(IROW,J)

• TRANS(IROW,J)

• ZMAX(IROW,J)

C. CODE EXECUTION

Polynomial coefficients describing the distribution of

the ratio of blade element chord to tip chord on a

projected plane.

Blade section definition parameter. If 0 the blade

element surfaces and centerline are described by

dk/ds = const. If not 0 they are defined by a fourth-

degree polynomial input on the next line of the data

set. See Reference 9, for a more complete

discussion.

Incidence angle array on streamlines.

Deviation angle array on streamlines.

Inlet/outlet segment turning rates on streamlines.

Transition point location on streamlines.

Maximum thickness point location on streamlines.

The name of the executable file running on the AA Department workstations, and

a 486 PC at the Turbopropulsion laboratory is Stream. To run the code on the

workstations simply type stream and enter. When running the code on the PC, you must

first activate the extended memory by typing os386 and enter, this is a requirement of the

FORTRAN compiler, and then up stream and enter. At this point you will be prompted

for the name of the input data file. This data file must be located in the same directory as

the executable file. If you stick to a prefix of less than seven alpha-numeric positions, and

a suffix of three letters, e.g. NASAR20.DAT, you will be O.K. After typing in the name of

the input file press enter and you will immediately be prompted for the name of the output

file. Use the same direction for the name of this file as was used for the input file. After

typing in the output file name press enter again. If the code has any trouble reading the

input file the program will stop at the point where the difficulty was encountered. By

45

checking the output file that, will be printed up until the time of the error, you should be

able to find the source of the error in the input.

Reminder:

An important note here is that a "dummy" editor, such as Notepad™ ,on the PC, or

jot, on the workstations should be used to create or modify the input file. Other

editors including DOS and word processing editors have been found to change the

formatted structure of the file, however slightly, enough to move the input data out

of position and prevent the correct input from being read.

Once the input is read the program will ask if you want an off design output file to

be created. At the time of this report the off-design code that was written to interface with

the design code was undergoing modifications. If yes or Y is selected the program will

create an input data set to the off-design code, however, given that the off-design program

is not up and running it would be prudent to type N. After the off-design prompt the

program will enter its aerodynamic and blade design (if activated) iteration. It is

recommended that the input parameter IBR be set to zero until the program is running to

take advantage of any error messages displayed during these iterations. After the programs

aerodynamic solution converges, you will be prompted to reset blade angles. This option

has not been exercised. It is up to the reader to investigate this "uncharted territory". If N

is entered at this prompt, the program will continue by stacking the blades one final time

and printing the coordinates. The output and, if selected, the off-design input file will be

placed in the same directory as the executable and the input files.

46

D. OUTPUT

The output of the code is broken up into four sections, consisting of the following;

1. A summary of the input.

2. Velocity diagrams and flow-field information at annular and blade edge

calculation stations.

3. Stage performance summary.

4. Fabrication coordinates if blade design option is activated.

Additionally, iteration output is provided, after the input summary, if the error message

parameter ffiR is set to zero. A copy of the output for the five runs of the code, as

described in part A of this appendix, is located in Reference 10.

The output of the code is formatted in a landscape orientation. Because of the

length of the file, up to more than eighty pages for a multistage compressor, the file must

be printed from a word processor on a PC. The conversion process involved in

transferring the file from DOS to the word processor tends to change the layout. Attempts

to print the correct layout of the output on a PC were unsuccessful. However, using the

following command,

>/usr/local/bin/a2ps.5.2 -p -nH -1 -1 -ns -nP -F8.5 FILENAME | Ip -dhp3si_l

on the AA Department workstations, it was possible to print the output. The current

operation of the networking system requires this exact input to specify all the parameters

needed for printing.

A final note concerning the output is that if the program is run using the PC

version, the output is broken up into two parts. The first part uses the name defined by the

user, and contains the performance information. The second part is located in a separate

file, using the same prefix and .COO for the suffix, and contains the blade fabrication

47

coordinate data. The fabrication coordinates of the output blade design are given for the

suction and pressure surfaces on planes parallel to the axis of rotation, relative to the

chord, and relative to the axial direction ("turbomachinery orientation"). When running the

program on the workstations these files are combined in the user defined output file.

Recall that the fabrication-coordinate data are only output when the blade design option is

activated.

E. CONCLUDING REMARKS

As noted previously, the AXIDES code is a 2-D code that produces the

aerodynamic and blade design (if desired), of single or multistage axial compressors. The

run time of the code is measured in seconds, and the output of the code is in a form that

can be readily used in mechanical analysis routines and in the fabrication process. The

AXIDES code is available commercially and can be conveniently used on a personal

computer. Although the code is based on 2-dimensional flow theory, if the application for

the compressor being designed does not require "cutting edge" design technology, the

short run time and conveniently formatted output make the AXIDES code a powerful tool

to use. It is understood that smaller companies that design compressors for use primarily

in an industrial capacity, and not for high performance aircraft, currently use the AXIDES

code as a final design tool. These companies can not make a major investment in the

development of their own design system and "reverse engineering" data base as do the

larger aircraft engine companies. Equally, they can not afford the engineering and

computer costs of using sophisticated CFD analysis codes.

48

CM

o> 0.500 0.500 0.5

.0000

HOCK

TABLE LE

.0

.0000 .0000 .0

0.0090

.0000

.0

.0000 .0000

0.00 0.00 0.0

8 . 80 8

. 80 6.7

0.50 0.50 0.5

.50 .50 .5

.0000

.0000

.0000

*•> ooo-o ooo

■o ■A OO OOOO «0 lA «A OOO

Q. E a

O U O 09 O

3

ä

0 0.500 0.500

00 0.5000

00 9.4000 .0100

00 .0100 .0100

TABLE SE TABLE

00 .0000 .0000

80 .0000 .0000

00 .0000 .0000

0.00 0.00

0.00

9.40 8.90

9. 10

0.50 0.50

0.50

.50 .50

.50

00 14.0000 .0150

00 18.0000 .0150

00 33.0000 .0150

8 iniA^O fOOOO— OlAintAtflOOO oo.-g-oooo-.oc, »•-

« 8 Hi o o"»-.-

°°| ~s °->z""> ill "*" K-

CO £ 355

1 0.

- 1.

S3 1 0

1.574 38.9700 1.0000

43996E-06

1 43579E-15

1 .6686 .7600 .8457

388.17 3BS.17 28

8. 17

10130.0 101)0.0 10110.0

.0000 .0000 .0000

-12.7000 -10.1600 -7.6300

7.6300 10.97)0 12.7000

2S.4000 25.4000 25.4000

25.1460 35

. 14

60 25. 14

60

-13.7000 -10.1600 -1.6200

7.6200 10.97)0 12.7000

t.5190 10.0960 10.5150

14.7320 15.2400 IS.2400

4000 .S000 .6000 .7000

4000 .5000 .6000 .7000

4000 .SOOO .6000 .1000

4000 .5000 .6000 .7000

4000 .5000 .6000 .7000

4000 .SOOO .6000 .7000

4000 .SOOO .6000 .7000

4000 .5000 .6000 .7000

4000 .5000 .6000 .7000

4000 .5000 .6000 .7000

4000 .5000 .6000 .7000

4000 .5000 ,6000 .7000

4000 .5000 .6000 .7000

4000 .SOOO .6000 .1000

4000 .5000 .«000 .7000

4000 .5000 .6000 .7000

4000 .5000 .6000 .7000

4000 .5000 .6000 .1000

4000 .5000 .6000 .7000

4000 .5000 .6000 .7000

4000 .5000 .6000 .7000

4000 .5000 .6000 .7000

-3.0

.05

.03

.0000

.0000

.0000

.0000

.0000

.0000

1.0000 7.

82 3 0

LE TABLE LE .0000

.0000 .0000 .0000

.0000 .0000 0.0000

.0000 .0000 o

■00

.00

.00

2.90 2.90 3.00

1-00

1 .0

00

1 .0

0

0.600 0.540 0.490

To TO Q. E a

»™S o 2 22 22 °° °° i*i'"- - o 2 £5 2- Zo -2 22S222222222222°oooooo

< OOOOOOO OlAO ' O O O l"* <■! r- O OlAO o OOOOOOO OOOO O O r*

• ** a o - "fUQ ü!"*1 • *• «««"•-'•'»«•-»«««««««««i'SSsS 3 ' ' " " - °

A

*•<!

•■ ign Of NAB

A 3

8 3 30 30 16100.0

0 .39S54990E-04

9 -.8199370SE-12

.3600 .4686

288.17 288.17

■BB. 17

lOUO.O 101)0.0

101)0.0

0 .0000 .0000

0 .0000

0 -25.4000 -30.3300

0 0.0000 4.0640

0 30.3300 34.0000

0 35.4000 35.4000

0 35.)900 35.1460

0 35.1460 35.1460

0 -35.4000 .30.3200

0 0.0000 4.0640

0 30.3200 14.0000

D 7.49)0 I.35S0

0 12.6660 13.9700

D 15.2400 15.2400

0243 .0312 .0406

0176 .0222 .0289

0132 .0163 .0210

0103 .0130 .0165

010) .01)0 .0165

010) .01)0 .01*5

010) .01)0 .0165

0103 .0130 .0165

0123 .0151 .0200

0140 .0182 .0343

0168 .0331 .0296

0373 .04)0 .0508

0320 ,0363 .042)

0383 .031) .0)60

0351 .0260 .0310

03)4 .0361 .039«

03)6 .0364 .0399

0241 .0269 .0)0«

0340 .0310 .0)17

0270 .0)0) .0341

0320 .0362 .0423

0)58 .0441 .0486

-17.0000 .0000

-2 1 .0000 .

0000

-13.0000 .0000

-7.0000 .0000

-1.0000 .00)0

1.0S00 .0070

.0100 .0100

ABLE SS TABLE TA

.0000 .0000

■0000 0.0000

.0000 0.0000

.00 .00

.00

2.90 2.60

1.10

0.50 0.75

1.000

0.6«0 0.640

0.300

s CO

to at *tt>q<

6 63070E*(

5OB40E-C

.2457

1.0000

288.

IT

IBB.11

101 JO

.0

101)0.0

.000

.000

-3.540

17.780

25.400

25.400

25.14«

-30.480

-2.540

17.780

7.102

11.747

15.240

019»

0141

0111

0089

0089

0089

0089

O0B9

0103

0110

0127

03)6

0290

02«)

02)9

0220

0222

022«

02)1

0348

0290

0)17

33.0000

21.0000

13.0000

-7.0000

-3.0000

2.050C

20.0000

AC 1.3000

.0000

.040«

.00

.00

4.40

6.20

«.35

1.000

0.360

* ,g

eg ^«--SsS2*S*SS:SS*2SX;«g£22£S2c:oSPS2-2:':"^*; s s;5„o.ooS" CL — ■ • «J m. o o -. i!C *C*C««7 — *~ 7 — *00.oooc'0°oooooooooooooooooooooc.oc 33333."-°°°°°::°°::

Z. Ä = Z Z O < < < < < a.

Figure 14. Examp.dat input data file

49

*.***..**.*.«***-*-****-*********** COLUMN **.*.***«***»» 11111111112222222222333333333 344444444445555555555666666666677777777778

12345678901234567890123456789012345678901234567890123456789012345678901234567890

TITLE ITG IBR

First Stage Redesign of NASA 2 Stage

NSTRM NROW NA NLOSS NTIP NHUB ROT

2 8 3 20 20 16100.0

CP(K) ,K=1,3

11

.23762070E+00

.81650840E-09

AR = 1-52

FLOW PRATIO

29.484 1.574

- .28462996E-06

.1257

.9257

288.17 288.17

10130.0 10130.0

.0000

.0000

-35.5600 -5.0800 15.2400

.2457 1.0000

288.17 288.17

10130.0 10130.0

.39556990E-04

CP(K) ,K=4,6

-.81993708E-12 .28442579E-15

CUMULATIVE FLOW FRACTION AT STREAMLINES

.3600 .4686 .5714 .6686

INLET TOTAL TEMPERATURE AT STREAMLINES

MOLWT SCALEF = = = = = ====== 28.9700 1.0000

LUNITI

1

LUNITO

1

288.17 288.17

288.17 288.17 288.17

.7600

288.17

INLET TOTAL PRESSURE AT STREAMLINES

10130.0 10130.0

.0000

.0000

INLET TANGENTIAL VELOCITY AT STREAMLINES

.0000 .0000 .0000

.8457

288.17

10130.0 10130.0 10130.0 10130.0 10130.0

.0000

.0000

AXIAL COORDINATE OF CASING WALL POINTS

-30.4800 -25.4000 -20.3200 -15.2400 -12.7000 -2.5400 0.0000 4.0640 5.0800 7.6200

17.7800 20.3200 24.0000

.0000

-10.1600 10.9730

.0000

-7.6200 12.7000

Figure 15. Examp.dat input data file (breakdown).

50

11111111112222222222333333333344444444445555555555666666666677777777778 12345678a01234567890123456789012345678901234567890123456789012345678S01234567890

RADIAL COORDINATE OF CASING WALL POINTS

25.4000 25.4000 25.4000 25.4000 25.4000 25 4000 25 4000 25.4000 25.4000 25.4000 25.3900 25.1460 25.1460 25.1460 25.1460 25.1460 25.1460 25.1460 25.1460 25.1460

AXIAL COORDINATE OF HUB WALL POINTS

-35.5600 -30.4800 -25.4000 -20.3200 -15.2400 -12.7000 -10.1600 -7.6200 -5.0800 -2.5400 0.0000 4.0640 5.0800 7.6200 10.9730 12.7000 15.2400 17.7800 20.3200 24.0000

RADIAL COORDINATE OF HUB WALL POINTS

7.3020 7.3020 7.4930 8.2550 9.1440 9.5890 10.0960 10.5150 11.0490 11.7470 12.6680 13.9700 14.3360 14.7320 15.2400 15.2400 15.2400 15.2400 15.2400 15.2400

LOSS PARAMETER ARRAY DIFFUSION FACTOR PARAMETER ARRAY

SET #1

•0167 .0199 .0243 .0312 .0406 .3000 .4000 .5000 .6000 .7000 .0123 .0143 .0176 .0222 .0289 .3000 .4000 .5000 .6000 .7000 ■0100 .0113 .0132 .0163 .0210 .3000 .4000 .5000 6000 .7000 .0080 .0089 .0103 .0130 .0165 .3000 .4000 .5000 6000 .7000 .0080 .0089 .0103 .0130 .0165 .3000 .4000 .5000 6000 .7000 .0080 .0089 .0103 .0130 .0165 .3000 .4000 .5000 6000 .7000 •0080 .0089 .0103 .0130 .0165 .3000 .4000 .5000 6000 .7000 .0080 .0089 .0103 .0130 .0165 .3000 .4000 .5000 6000 .7000 .0090 .0103 .0122 .0153 .0200 .3000 .4000 .5000 6000 .7000 .0092 .0110 .0140 .0182 .0243 .3000 .4000 .5000 6000 .7000 .0104 .0127 .0168 .0221 .0296 .3000 .4000 .5000 .6000 .7000

SET #2 ======

.0309 .0336 .0373 .0430 .0508 .3000 .4000 .5000 6000 . 7000

.0272 .0290 .0320 .0362 .0423 .3000 .4000 .5000 6000 .7000

.0250 .0263 .0282 .0313 .0360 .3000 .4000 .5000 6000 .7000

.0230 .0239 .0253 .0280 .0310 .3000 .4000 .5000 6000 .7000

.0211 .0220 .0234 .0261 .0296 .3000 .4000 .5000 6000 .7000

.0212 .0222 .0236 .0264 .0299 .3000 .4000 .5000 6000 .7000

.0214 .0226 .0241 .0269 .0306 .3000 .4000 .5000 6000 .7000

.0218 .0231 .0248 .0278 .0317 .3000 .4000 .5000 .6000 .7000

Figure 16. Examp.dat input breakdown (cont'd).

51

*•..**•****»••**********♦********** COLUMN ***•«*♦♦***♦«»« 11111111112222222222333333333344444444445555555555666666666677777777778

1234567890123456789012345678901234567890123456789012345678901234 5678901234567890

SET «2 (CONT'D)

0233 0248 . 0270

0272 .0290 .0320

0294 .0317 .0358

.0

.0

.0

.0

.0

.0

.0

.0

.0

.0

.0

.0303 .0347 .3000 .4000

.0362 .0423 .3000 .4000

.0441 .0486 .3000 .4000

SET «3

.5000

. 5000

. 5000

CALCULATION STATION AND BLADE ROW DATA

.6000 .7000

.6000 .7000

.6000 .7000

•2.0

.05

.03

AA ZTIP ZHUB BT BH BLEED

ANNULAR ANNULAR ANNULAR ANNULAR ANNULAR

-32.0000 -21.0000 -13.0000 -7.0000 -3.0000

-32.0000 -21.0000 -13.0000 -7.0000 -3.0000

.0000

.0000

.0000

.0000

.0030

.0000

.0000

.0000

.0020

.0030

.0000

.0000

.0000

.0000

.0000

AA ZTIP ZHUB BT(LE) BH(LE) BLEED(LE)

ROTOR 2.0500 2.0500 .0070 .0070 .0000

DLIM ALIM BT(TE) BH(TE) BLEED(TE) CRENGY BMATL NXCUT

.4600 -20.0000 .0100 .0100 .0000 1.0000 7.823 0

ILOSS

1

OP AA AB BB CC DD EE EB

COORD TABLE SS TABLE TABLE TABLE TABLE LE

CHOKE

.0000

Figure 17. Examp.dat input breakdown (cont'd).

52

«***...**.*♦*..*.,»*«..»».. COLUMN ****•****»•♦**.♦**•**•*.***♦**.,,.*

11111111112222222222333333333344444444445555555555666666666677777777778 12345678901234567890123456789012345678901234567890123456789012345678901234567890

# BLADES SOLIDITY TILT

44.0000 1.3000 .0000

PRA

.0000

PRB

.0000

PRC

.0000

PRD

. 0000

PRE

.0000

TALE

.0055

TAMAX

.0327

.00

.00

5.40 3.60

0.00 1.000

0.700

0.420

TBLE

.0000

TBMAX

.0496

.00

.00

4 .40 6.20

0.25 1.000

TCLE

.0000

TCMAX

.0000

TDLE

0.0000

TDMAX

0.0000

TÄTE

.0055

CHORDA

.0000

TBTE

.0000

CHORDB

.0000

INCIDENCE ANGLE ARRAY ON STREAMLINES

■00 .00 .00 .00 .00

DEVIATION ANGLE ARRAY ON STREAMLINES

2-90 2.60 2.80 2.90 7.80

TCTE

.0000

CHORDC

.0000

.00

2.90

INLET/OUTLET SEGMENT TURNING RATE ON STREAMLINES

°-75 1.00 1.00 1.000 0.50 1.000

0.680 0.360

TRANSITION POINT LOCATION ON STREAMLINES

0.640 0.620 0.600 0.540 0.660 0.300

MAXIMUM THICKNESS POINT LOCATION ON STREAMLINES

0.500 0.500 0.500 0.500 0.500 0.500 0.500 0.5000 0.5000 0.5000

AA

STATOR

ZTIP

9.4000

ZHUB

9.4000

BT(LE)

.0100

BH(LE)

.0100

BLEED(LE)

.0000

TDTE

0.0000

IDEF

0

.00

3.00

1.00

0.480

0.500

Figure 18. Examp.dat input breakdown (cont'd).

53

.*.....*.........* .**.**********"* COLUMN .-..*♦****-************-**'** 11111111112222222222333333333344444444445555555555666666666677777777778

12345678901234567890123456789012345678901234567890123456789012345678901234567890

DLIM ALIM BT(TE) BH(TE) BLEED(TE)

.7000 1.0000 .0100 .0100 .0000

ILOSS OP AA AB BB CC DD EE EB

2 COORD TABLE SS TABLE TABLE SHOCK TABLE LE

# BLADES

48.0000

SOLIDITY

1.2700

TILT

.0000

PRA

.0000

PRB

.0000

PRC

.0000

PRD

.0000

TALE TBLE TCLE TDLE TÄTE TBTE TCTE

.0061 0.0080 .0000 .0000 .0061 0.0000 .0000

TAMAX

.0667

TBMAX

0.0200

TCMAX

.0000

TDMAX

. 0000

CHORDA

.0000

CHORDB

.0000

CHORDC

.0000

INCIDENCE ANGLE ARRAY ON STREAMLINES

0.00 0.00 0. 00 0.00 0.00 0.00 0.00

0.00 0.00 0.00

11.10

DEVIATION ANGLE ARRAY ON STREAMLINES

14.20 9.40 8.90 8.80 8.80 8.80

8.70 8.80 9 . 10

0.50 0.50

.50

.50

AA

ANNULAR ANNULAR ANNULAR

0.50 0.50

.50

.50

ZTIP

14.0000 18.0000 23.0000

INLET/OUTLET SEGMENT TURNING RATE ON STREAMLINES

0.50 0.50 0.50 0.50 0.50

0.50

TRANSITION POINT LOCATION ON STREAMLINES

.50 .50 .50 .50 .50 .50

ZHUB

14.0000 18.0000 23.0000

BT

.0150

.0150

.0150

BH

.0150

.0150

.0150

BLEED

.0000

.0000

.0000

NXCUT

0

CHOKE

.0000

PRE

.0000

TDTE

.0000

IDEF

0

0.00

8.70

0.50

.50

Figure 19. Examp.dat input breakdown (cont'd).

54

Option

Input Data Field for the Program AXIDES (General or Overall Data)

|T| ,

|y|.. Column *•"••• • |

I Pi 111111111 122222222223333333333444444444455555555555666666666777777777781

|e 1123456789012345678901234567890123456789012345678901234567890123456789012345676501

IF NSTRH>9

If NTIP>8

IF «TIP>16 IF HTIP>24

IF NTIP>32

IF miP>8

IF MTIP>16

IF NT1P>2«

IF NTIP>32

IF nmn>6 IF NHU8>16

IF HHU8>2<

IF NHUS>32

IF «HUB>8 IF NHUB>16

IF «KUB>24

IF MHUB>32

IF NLDSS>0. J=1

J*2

J=NSTRM

IF KLOSSH, J«1

J«2

J-HStHH IF ltLOSS>2, J-1

J>2

IF «LOSS>3, J«1

J=NSTRN IF «L0SS>4, J«1

J=2

J=KSIIW |

l = l = |6|

M

M M |B|

|6|

M l = l l«l M l=l |6|

W IM |s| |e| M |e| |6|

|e| |6|

|6|

|B|

|s| l=l |B| |6|

|6|

|6| M |6| |6|

M |6| |S|

■ «•I |6|

M |S|

|S| M |e| |e| |6| |e| M l=l |6|

<■!

«■I

TITLECK), K=1,18

HSTRN mOU tl» HLOSS «TIP PJHUB ROT

(CP(K),K>1,3)

(CP(IC),K=4.6)

(Cumulatntive Floy Free. From Tip a

(Inlet Totat Tenp. 8 Streamlines)

(Inlet Total Press. S Streamlines)

(Inlet Tang. Vel. 9 Streamlines)

(Aitial Coord, of Casing Wall Pts.)

(Radial Coord, of Casing wall Pts.)

(Axial Coord, of Hub Us 11 Pts.)

(Radial Coord, of Hub Wall Pts.)

FLOW

Streamlines)

PR41I0 HOIWT

LUNIT0|

(FLOCK).K=2,NSTRM) |

(FL0OC),IC = 10,HSTRH)|

(Loss Parameter)

(Lost Parameter)

(Loss Parameter)

(Loss Parameter)

(Loss Parameter)

(DLOSU (DLOSOC

(DLOSOC CDlOSCt

(DLOSOC

(DLOSOC

(DLOSOC

(DLOSOC

(DLOSOC (DLOSOC, (DLOSOC,

(DLOS(K,

(DLOSOC, (DLOSOC, (DLOS(K,

(DLOS(K

(DLOSOC

(DLOSOC

(DLOS(«C,

(DLOSOC,

(DLOSOC,

(DLOSOC,

(DLOSOC.

(DLOSOC,

(DLOSOC

J.D.K .J.D.X

J.D,*' .J.D.K'

. J, 1).K>

,J.2>.K<

J,2),K'

J,2).IC'

J.2),»C"

J.2)."C'

J.3>.K-

j.3),r>

j,3).r.

J,3),«-

J.3>,K-

J.*).r« J.*>.ic«

J,4),C«

J. «>.<:= J,4),r«

J.5),»c=

J,5),IC>

J.Sj.K'

J.5>,K'

= ',5)

1.5)

=1.5)

1.5)

■1.5) ■1.5)

■1.5) ■1.5) ■1.5) ■1.5) ■1.5)

■1.5) =1.5)

■1.5) ■1.5)

1.5)

1.5)

'1.5)

1.5)

1.5)

1.5)

1.5)

1.5)

1.5)

1.5)

ITC IHR I 18A4, 214

SCALEF | 615, 5F10.4

LUHITI | 3E20.8, 10X, 110

3E20.8, 10X, 110

8F10.4

8F10.4

| 8F10.4

| 8F10.4

| 8F10.4

| 8F10.4

| 8F10.4

| 8F10.4

| 8F10.4

| 8F10.4

j 8F10.4

I 8F10.4

| 8F10.4

j 8F10.«

j 8F10.4

j 8F10.4

j 8F10.4

| 8F10.4

j 6F10.4

j 8F10.4

| 8F10.4

| 8F10.4

j 8F10.4

| 8F10.4

| 8FI0.4

j SF10.4

j 8F10.4

| 8F10.4

(0FTABOC.J,1),tC=1,S>| 10F8.4

(DFTABOC,J,1),K=1,5)| 10F8.4

(DFTABCK,J,1>,K=1,5>| 10F8.4

(DFTABOC,J,1),r=1,5>| 10F8.4

(DFTABOC,J,1),IC=1,5)| 10F8.4

(Dfactor Array) (0FTABOC,J,2),K=1,5)| 10F8.4

(DFTABOC,J,2),K=1,5)j 1QF8.4

(DFTABOC,J,2),IC=1,S)| 10F8.4

(DFTABOC,J,2),«C=1,5)j WFB.*

(DFTABOC,J,2),IC=1,5)| 10F8.4

(Dfactor Array) (DFTABOC,J.3>,r*1,5)j 10F8.4

(DFTAB(rC,J,3),r=1,5)| 10F8.4

(DFTABOC,J,3),K=1,5>| 10F8.«

(0FTAB(r:,.J,3),<C=1,5)| 10F8.4

(DFTABOC, J, 3), BC-1.5>| 10F8.4

(Dfactor Array) (DFTABOC,J,*),r=1,5)| 10F8.«

(DFTABOC,J,4),FC=1,5)| 10F8.4

(PFTABOC.J,4),11=1,5)1 10F8.4

(DFTABOC,J,4),r=1,5)| 10F8.4

(0FTAB(«C,J,4),IC=1.5>| I0F8.4

(Dfactor Array) (DFTABOC,J,5),K=1,5)j 10F8.4

(DFTABOC,J,5),IC=1,5)| I0FB.4

(DFTABOC,J,S),IC=1,5)| 10F8.4

(DFTABOC, J, 5), IC=1,5)| 10F8.4

(DFTABOC,J,5),tC=l,5)| 10F8.4

(T0(1,J),J=1,HSTPH)

(T0(1,J),J=9,HST°H)

(P0(1,J),J=1,NSTRH)

<P0(1,J),J=9,HSTSH)

(VTH(1,J),J=1,NS1RH)

(VTH(1,J),J=9,NSTRM)

(KTIP(I),1=1,HTIP)

(XTIP(I),I=9,HTIP>

fKTIPCI>.I'17.HTIP)

(XTIP(I),I=25,HT!P)

(XTIP(I),I=33,NTIP)

(FtTIP<l),l=1,NTIP)

(RT1P(I),I=9,MIP)

(RTIP(I),I»17,HIIP)

OtTIP(l),I=25,HTIP)

(RTIP(1),I=33,HTIP)

(XHU8(I),I=1,HTIP)

(XHUB(I),I=?,HTIP)

(XHUB(I),I«17,HTIP)

(XHUB(I),I=25,HTIP)

(XHUB(t),1=33,»TIP)

(RHUB(I),I=1,HT1P)

(RHUB(I),I=9,HTIP)

(RHU8(I),I=17,NTIP)

(RHU8(I),I=25,HT1P)

(«HUB(I),I=33,HTIP)

(Ofactor Array)

Figure 20. General input data field format.

55

Tnput Data Field for the Proaram_AXIDES (Data for Annular Stations & Rotors)

I'l 1VI Inl

11111111112222222222333333333344444444445555555555566666064677777777778 Format

I'l 12345678901234567890123456789012345678901 234567890123456789012345678901234567890

1*1 ZT1PO) ZHUBO) BT(1) 8H(1) BLEED(I)

1*1 ZT1PC2) ZHU8C2) BT(2) BH(2) BLEED(2) A4.6X.5F10.4

1*1 • A Z1IP13) ZHUB(3) BT(3) BH(3) BIEED(3) A4.6X.5F10.4

1*1 AA ZTIP(4) 2HU8<4) BI(4) BH(4) BLEED(4) A4,6X,5F10.4

l»l At ZTIP(IXAB) ZHUB(IHAB) Bill) BH(1) BLEEO(l) A4.6X.5F10.4

|R| DLIHC1RU) ALIMURU) Bl(l) BH( 1 ) BLEEO(I) CRENGr(IR)BMATLUR) NXCUT(IRU) 7F10.7,110

|R| I LOSS OP« OP OPO AA AB BB CC OD EE EB CKOKE(IRU) l5,1X,2A4,2x,3(A4,

|R| BLAOESCIRU) SOLID(IRU) TILKIRU) PRACIRU PRB(IRU) PRC(IRU) PRD(IRU) PRE(IRU) 8F10.4 6F10.4

If UOSS(1ROU)>*0|R PTT(IRU) (PTC(J.IRU) ,J=1.5>

|B TALE(IRU) TBLE(IRU) ICLE(IRU) TOLE(IRU) TATEURU) TBTE(IRU) TCTE(IRU) TDIE(IRU)

|R TAMAXC ) IBHAXC ) TCKAXC ) TOHAXC ) CHORDA( ) CHORDB! ) CHOROCC ) IDEF(IRU) 7F10.4.110

IF IDEFC1ROU)#0 lR (ACF(J,IROU),J=1.4> (BCF(J,1ROU),J=1,4> 8F10.6

IF lt>EF(IROU>#0 1« <CCF<J,IROU),J=1,4> (0CF(J.IROU),J=1,4)

IF IDEF(IROU>#0 1» <ACR(J,1R0U).J=1.4) (BCS(J.IROU),J=1,4) 8F10.6

IF IDEF(IROU>*0 1« (CCRCJ.IROU>,J=1,4) (DCR(J,lROU),J=1,4)

IF IDEFUROWjfO 1" (ELE(J,IROU),J=1.'> (ETE(J,1R0U),J=1,4) 8F10.6

8F10.6 IF IDEF(1ROU)#0 1« <ATF(J.mOU),J=1,4) (8TF(J,IROU),J=1,4>

IF IDEF(IROU>*0 1« (CTF(J,IROU),J=l,4> (DlF(J,IROU),J=1,4)

IF 10EF(IR0U)iD 1" <ATR(J,IROU),J=1,«) (BTR(J,IROU),J=l,4) 8F10.6

IF IDEF(IROU)#O 1« <CTR(J,IROU>,J=1,4> (DTR(J,IROU),J=1,4)

;F AA=TABIE 1" (Incidence Angle Array on Streamlines) (1KC(IR0U,J),J=1,«STRM)

IF NSTRM>8 1« (IRC(IROU,J),J=9,tlSTRK) 8F10.4

IF BB*TABLE 1« (Deviation Angle Array on Streamlines ) (DEV(IROU,J),J=l,WSTRM) 8F10.4

IF NSTRH>8 1« (DEV(1ROU,J),J*9,NSTRM) 8FI0.4

IF COTABLE 1« (Inlet/Outlet Segment Turnina Rate) (PHKIROU,J>,J = 1,HSTRM) 8FI0.4

IF HSIRtOB 1« (PHI(IROU,J),J=9,KSTRH) 8F10.4

IF DATABLE 1« (Trans. Point Location) (TRAHS(IROU,J),J=l,HSIRH) 8FI0.4

IF «S7«M>8 1« (TRANS(IROU,J),J=9,NSTRH) 8F10.4

IF EE-TABLE 1« (Max. Thickness Point Locat on) (ZMAXIISOU,J),J-1.HSTRM) 8F10.4

IF «ST«M>8 1" (ZMAX(IROU.J),J»9,IISTRt<) 8F10.«

IF 0P*VEL.01A. 1« (ZTEMP<1-1,J),J=1,5> (RTEHP(I-I,J),J«1,5) 10F8.4

IF OP'VEL.DIA. 1« (ZTEMP(1,J),J=1,5) (R1EHP(I,J),J=1,5) 10F8.4

Figure 21. General input data field (cont'd).

56

Input Data Field for the Program AXIDES (Data for Stators & Annular Stations)

M**"* Colljm OP' i0° |P| 1111111111222222222233333333334444«44444555555555556©666466677777777778|

|e|1234567890123456789012345678901234567890123456789012345678901234S678901234567890|

IF 110SS(IROU>>

IF 1DEF(1R0U)#G

IF IDEF(IROU)«

IF IDEF(1R0U)#C

IF ID£F(IROU)#0

IF IDEF(IROU)tO

IF lDEF(IROU)«

IF IDEF(1ROU>«0

IF IOEF(I«OU)«0

IF IDEF(mOU)ff)

IF AA=TA8LE

IF NST«K>B

IF IB-TABLE

IF MSt«M>«

IF COTABLE

IF HSTRM>8

IF DD*TA8LE

IF »STRM>8

IF EEMA8LE

IF KSTRM>8

IF OP*VEL.DIA.

IF OP'VEL.OIA.

|S|

1*1 I* 1*1

'°l*l 1*1 1*1 1=1 1*1 |s| 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1 1*1

AA 2TIF>(INAB> ZHUBOHAB) BI(I)

DLIM(IHU) ALIM(IRU) 81(1) BH(I)

I LOSS OPM Of OPO AA AB SB

BIADES(IRW) SOLIO(IRU) TILT(IRW) PRA(IRU)

PTTURU) (PTC(J,IRV),J=1,5>

lALE(IRU) IBLE(IRW) TCLE(IRU) TOLE(IRU)

TAKAX( ) 7BHAX( ) ICMAXt ) TDHAX( )

(ACF(J,IR0U),J=1,4)

(CCF(J.IR0W),J=1,«)

(ACR(J,IR0U),J=1,4)

(CCR(J,IR0W),J=1,4)

(ELE(J,IR0U),J=1,4)

(ATF(J,IR0U),J-1,4)

(CTF(J,IR0U),J=1,4)

(ATR(J,IR0U),J=1,4)

(CTR(J,IR0U),J=1,4)

(Incidence Angle Array on Streamlines)

(Deviation Angle Array on Streamlines)

(Inlet/Outlet Segment Turning Rate)

(Trans. Point Location)

(Max. Thickness Point Location)

BH(|) BLEEO(I)

BLEEO(I)

CC OD

PRB(IRW) PRC(IRU)

NXCUK

EE EB CHOICE ( PRD(IRW) PRE(

(ZTENP(I-1,J),J*1,S>

(ZTEMP(I,J),J=1,5)

TA1ECIRU) TBTE(IRU) TCTE(IRU) TDTE(

CHOROAC ) CHOROBC ) CHOR0C( ) IDEF(

(BCF(J,IR0U),J=1,4>

(DCF(J,IR0U),J=1,4>

(BCR(J,IR0U),J=1,4>

(DCR(J,IR0U),J=1,4)

(ETE(J,1R0U),J=1,4)

(BTF(J,1R0U),J-1,4)

(0TF(J,IROU),J=1,4)

(BTR(J,IR0U),J«1,4)

(DTR(J,IR0U),J=1,4)

(INC(IROV,J),J-l,IISTRM)

tmC(IR0U,J),J=9,IISTRM)

(0EV(IROU,J),J=1,RSTRM)

(DEV(IR0V,J),J-9,»STRH)

(PHI(IR0W,J),J=:1,NSTRM)

(PHI(IR0U,J),J>9,NSTRM>

(TRANS(IR0U,J),J*1,IISTRH)

(TRAHS(1R0U,J),J=9,NSTRH)

(ZMAX(IR0U,J),J<1,NSTRM)

(2KAX(1R0W,J),J«9,IISTRH)

(RTEMP(l-1,J),Jr1,5)

(RTEMP(l,J),Jc1,5)

I A4.6X.5F10.4

IRU)| 7F10.7,I10

IRU)| I5,1X,2A4,2*.3<A4,

IRU)| 8F10.3

| 6F10.4

IRU)| 8F10.4

IRU)| 7F10.4.I10

| 8F10.6

| 8F10.6

| 8F10.6

| 8F10.6

| 8F1Ö.6

j 8F10.6

| 8F10.6

| 8F10.6

j 8F10.6

| 8F10.4

| 8F10.4

j 8F10.4

j 8F10.4

j 8F10.4

j 8F10.4

| BF10.4

j 8F10.4

| 8F10.4

j 8F10.4

| 10F8.4

10F8.4

l-l l-l l-l l-l |A| AA

|A| AA

|A| AA

2T1P(I) IKUB(I) IT(I) IR(I> BLEEO(I) 2TIP(I) 2HUB(I) BT(I) BH(I) BLEED(I) 2TIP(I) ZHUB(I) BT(I) BR(I) BLEEO(I)

| A4,6X,5FI0.4 j A4.6X.5F10.4 j A4.6X.5F10.4

IF NXCUT(IROW><0 |C|

IF NC>8 |C|

IF N016 |C|

l-l l-l l-l l-l

(Radial Location of Blade Sections) (XCUT(J),J-1,NC>

(XCUT(J),J'9,NC)

(XCUT(J),J-17,NC)

| 8F10.4

j 8F10.4

j 8F10.4

Figure 22. General input data field (cont'd).

57

58

APPENDIX B. AXIDES OUTPUT

This appendix contains the output file Examp.out, associated with input file

Examp.dat. The input file is the example provided with the code when it was purchased

and was the first to be successfully run on the AA Department workstations.

59

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92

APPENDIX C. TEST RIG

This appendix contains information pertaining to the transonic axial compressor

test rig located at the NPS TPL. Included is the general procedure involved in operating

the test rig, and its disassembly and reassembly. This information is provided to fill the

void that currently exists in the documentation associated with maintenance and operation

of the test rig.

A. OPERATING PROCEDURE

A photograph of the control panel, from which the operation of the test rig is

controlled and monitored, is located in Figure 23. The panel is used to control the supply

of air from the laboratory's 12 stage compressor to the test-rig drive turbine, to relieve the

axial load on the drive shaft bearings by porting low pressure air to a balance piston inside

the compressor hub, and to remotely operate the hydraulic throttle located in the

compressor-intake housing. It is also used to monitor the operating temperatures of the

drive-shaft bearings, the RPM, and levels of vibrations. The following is a description of

the operating procedure.

1. The technician starts and warms up the laboratory air-supply compressor (1

hour), with the bypass (dump) valves in the open position.

2. Turn on the control-panel activation switch (located beneath the control-panel

counter-top).

3. Ensure that air is being supplied to the balance piston, and bearing oil-mist

cooling lines. This air is supplied through the shop air lines, by the Elliot

compressor.

4. Set the oil-mist cooling, air-oil regulators, to 35 psi, and 12 drops of oil per

minute (count the number of drips over a three minute interval and average).

5. Check that the bearing temperature readings are indicating test-cell ambient

values (not wandering).

93

6. Close supply-air dump valves (2 ea.),until a two-to-one pressure ratio across

the laboratory air supply compressor is indicated (on gauge to the right of

control panel), and the noise level in the control room is acceptable (some

combinations of valve settings are noisy).

7. Open air supply valve and close dump valves partially until desired RPM is

indicated, while maintaining the two-to-one pressure ratio across the air

supply compressor.

8. As the RPM increases, the air supplied to the balance piston must be increased

to compensate for the axial load on the bearings. If the value displayed on the

analogue readout is maintained at the level indicated before start-up (with no

air supplied to the balance piston), the axial load produced by the spinning

rotor will be negated by the pressure on the balance piston. This will result in

approximately zero axial load on the high speed bearings.

9. Set throttle valve as required, and correct air-supply valve to maintain RPM.

During operation of the test rig, RPM, bearing temperature, balance piston, and

vibration indications must be monitored continuously. Additionally, periodic checks of the

oil-mist regulators is recommended.

B. DISASSEMBLY AND REASSEMBLY

The disassembly and reassembly performed in association with this report was

documented on video tape. This tape will be maintained with the records of the test rig for

guidance in future overhaul efforts. A few notes, however, concerning the disassembly and

reassembly of the machine should be emphasized.

94

Disassembly

• The removal of the inlet nozzle bell from the hub housing is tricky. Extreme care must

be taken to prevent the shroud (attached to, and removed with, the nozzle bell), from

coming into contact with the rotor blade tips. To remove the nozzle bell and shroud

assembly without incident, pull out the positioning pins, and un-screw the bolts

securing the assembly to its support structure. Then attach the test-cell overhead crane

to the eye bolt on the top of the nozzle and raise the crane until there is a slight tension

on the chain. Attach a winch to the nozzle-inlet face, and the piping from the inlet

housing (outside the test cell), and gently pull it axially away from the hub housing. If

the winch cables are evenly distributed around the nozzle face the assembly should

slide along the channels fastened to the nozzle-bell support structure that ensure

clearance of the rotor blade tips. The overhead crane will prevent the assembly from

dropping as it clears the edge of the nozzle base.

Reassembly

► Ensure that markings on splines are located during disassembly and aligned during

reassembly. These markings line up the high speed rotational components as they were

when the machine was originally balanced.

' The pre-load on the high speed bearings is defined as "light". A mere 2-3 in.-lb.

torque, on the rotor mounting bolt, is all that is desired to keep the bearings, rotor and rotor shaft in place.

Perform a new calibration of the torque balance each time the machine is reassembled.

Check bearing oil-mist cooling lines, bearing temperatures, and internal

instrumentation lines for proper operation/indications prior to reassembling.

95

C. INSTRUMENTATION

The test runs and overhaul of the test rig, as described in this report, have been

primarily driven by the desire to return the machine to routine operation. The myriad of

instrumentation, that was added in a piece-meal fashion over the fifteen years of operation

of the rig, is in need of overhaul. Currently the data-acquisition system at the TPL is in

transition. The Hewlett-Packard Basic workstation and programs that have been used for

data acquisition are being augmented and possibly replaced with a PC data acquisition

controller using Labview™ A VXI-Bus mainframe has been purchased, and obsolete

scanners are to be replaced. All instrumentation lines from the test-cell to the data

acquisition area must yet be verified or replaced before performance mapping of the new

stage can begin.

96

Figure 23. Transonic Compressor Test Rig Control Panel.

97

98

LIST OF REFERENCES

1. Paige, G. C., Measurement of Case Wall Pressure Signatures in a Transonic Compressor Using Real-Time Digital Instrumentation. M.S. Thesis, Naval Post Graduate School, June 1976.

2. West, J.C., Digital Programmable Timing Device for Fast Response Instrumentation in Rotating Machines., M.S. Thesis, Naval Postgraduate School, December 1976.

3. Simmons, J. M., Data Acquisition and Analysis Techniques for Measuremen of Unsteady Wall Pressures in a Transonic Compressor. M.S. Thesis, Naval Postgraduate School, July 1977.

4. Dodge, F.J., Development of a Temperature-Pneumatic Probe and Application at the Rotor Exit in a Transonic Compressor M.S. Thesis, Naval Postgraduate School, June 1976.

5. Hawkins, W.R., Determination of the Blade-Element Performance of a Small Transonic Rotor, M.S. Thesis, Naval Postgraduate School, December 1976.

6. Erwin, J. R., A Review of the Design of the NPS/TPL Transonic Compressor Contractor Report NPS67-83-004CR, Naval Postgraduate School, Monterey' California, March 1983.

7. Neuhoff, F., Shreeve, R.P., and, Fottner, L., Evaluation of the Blade-to-Blade Flow From a High Speed Compressor Rntnr ASME Paper 86-GT-117, Presented at the International Gas Turbine Conference and Exhibit Düsseldorf West Germany, June 8-12, 1986.

8. Sanger, N.L., Design of a Low Aspect Ratio Transonic Compressor Stag* TTcinp CFD Techniques, ASME Paper No. 94-GT-236, Presented at the International Gas Turbine and Aeroengine Congress and Exposition, The Hague The Netherlands, June 13-16, 1994.

9. Crouse, J.E., and Gorrell, W.T., Computer Program for Aerodynamic and Binding Design of Multistage Axial-Flow Compressor NASA Technical Paper 1946, NASA Lewis Research Center, Cleveland, Ohio, December 1981.

99

10. Reid, W.D., AXTDES Code Input and Output Files. Turbopropulsion Laboratory Technical Note, TPL-TN-95- 01, Naval Postgraduate School, Monterey, California, September 1995.

11. Denton, J.D., An Improved Time-Marching Method for Turbomachinery Flow Calculation, ASME Paper 82-GT-239, Presented at the 27th International Gas Turbine Conference and Exhibit, London, England, April 18-22, 1982.

12. McNally, W.D., FORTRAN Program for Calculating Compressible Laminar and Turbulent Boundary Layers in Arbitrary Pressure Gradients. NASA TN D-5861, NASA Lewis Research Center, Cleveland, Ohio, 1970.

13. Denton, J.D., The Use of a Distributed Body Force to Simulate Viscous Effects in 3D Calculations. ASME Paper 86-GT-144, Presented at the International Gas Turbine Conference and Exhibit, Düsseldorf, West Germany, June 8-12, 1986.

14. Katsanis, T., FORTRAN Program for Calculating Transonic Velocities on a Blade-to-Blade Stream Surface of a Turbomachine. NASA TN D-5427, NASA Lewis Research Center, Cleveland, Ohio, 1969.

15. Katsanis, T., and, McNally, W.D., Revised FORTRAN Program for Calculating Velocities and Streamlines on the Hub-Shroud Midchannel Stream Surface of an Axial-, Radial-, or Mixed-Flow Turbomachine or Annular Duct. NASA TN D- 8430, NASA Lewis Research Center, Cleveland, Ohio, March 1977.

16. National Aeronautics and Space Administration Report No. N65-23345, Aerodynamic Design of Axial-Flow Compressors. NASA SP-36, Washington, DC, 1965

17. Shreeve. R.P.. Report on the Testing of a Hybrid (Radial-to-Axiall Compressor. Report No. NPS-57Sf73112A, Naval Postgraduate School, Monterey, California, November 1973.

100

INITIAL DISTRIBUTION LIST

1. Defense Technical Information Center 2 Cameron Station Alexandria, VA 22304-6145

2. Dudley Knox Library 2 Code 052 Naval Postgraduate School Monterey, CA 93943-5101

3. Chairman 1 Department of Aeronautics and Astronautics CodeAA Naval Postgraduate School 699 Dyer Road - Room 137 Monterey, CA 93943-5106

4. Professor R.P. Shreeve 10 Department of Aeronautics and Astronautics Code AA/SF Naval Postgraduate School 699 Dyer Road - Room 137 Monterey, CA 93943-5106

5. Professor G.V. Hobson 1 Department of Aeronautics and Astronautics Code AA/SF Naval Postgraduate School 699 Dyer Road - Room 137 Monterey, CA 93943-5106

6. Commander I Naval Ar Systems Command CodeATR4.4.T 1421 Jefferson Davis Highway Arlington, Virginia 22243

101

7. Naval Air Warfare Center Aircraft Division Code AIR 4.4.3.1 [S. McAdams] Propulsion and Power Engineering, Bldg. 106 Patuxent River, Maryland 20670-5304

8. Curricular Officer, Code 31 Naval Postgraduate School Monterey, CA 93943-5002

9. Mr. Thomas J. Reid 33 Crestwood Street Syosset, NY 11791

10. Mr. James Crouse No. 7 ST. Mary's Allegany, NY 14706-9672

11. Mr. Nelson Sänger 752 Elmwood Road Rocky River, OH 44116

102