Static pressure contours of turbine rotor blade tips at high Mach number

Bailey, E. E.

Miller, G. R.

English

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1. Report No. 2. Government Accession No. 
NASA TM X-2170 
4. Title ... d Subtitle 
STATIC-PRESSURE CONTOURS IN THE BLADE PASSAGE AT 
THE TIP OF SEVERAL HIGH MACH NUMBER ROTOFE 
3. Recipient's Catalog No. 
5. Report Date 
February 197 1 
6. Performing Organization Code 
7. AuthorW 
Genevieve R. Miller and Everett E. Bailey 
9. Performhg Organization Name and Address 
Lewis Research Center 
National Aeronautics and Space Administration 
16. Abstract 
Static pressure  contours across  the blade passage a t  the tip of three high Mach number rotors  are 
presented. Inlet relative Mach numbers vary from 1.23 to 1.59. Flow rates vary l rom near  
choke values to  those near blade stall. Blade shapes include both double c i rcular  arc and multiple 
circular arc. The s ta t ic  pressure coniour plots indicate the location of the passage shock waves 
over a range of Mach numbers and operating conditions. Qualitative comparisons of flow condi- 
tions through blade passages for the two blade shapes can be made. However, a quantitative 
analysis c;f the data is not attempted. 
- 
8. Performing Organization Report h,. 
E-5875 
10. Work Unit No. 
720-03 
11. Contract or Grant No. 
Cleveland. Ohio 44135 
12. Sponsoring Agency Name and Address 
National Aeronautics and Space Administration 
Washington, D. C. 20546 
18. 3istribution Statement 
Unclassified - unlimited 
13. Type of Report and Period Covered 
Technical Memorandum 
14. Sponsoring Agency Cod? 
this page) I 21. No. of Pages I 22. Rice' 19. Security Classif. (of this report) 20. Security Classif. ( 
Unclassified Unclassified 17 $3.00 
STATIC-PRESSURE CONTOURS I N  THE BLADE PASSAGE AT THE 
TIP O F  SEVERAL HIGH MACH NUMBER ROTORS 
by Genevieve R. Miller and Everett E. Bailey 
Lewis Research Center 
SUMMARY 
Static pressure contours across  the blade passage at the t ip of three high Mach nuin- 
ber rotors were made for inlet relative Mach numbers varying from 1.23  to  1. 59. 
rates for these contours vary from values near choke to those nea: blade stall. Blade 
shapes are both double circular arc and multiple circular arc.  
The static pressure contour plots indicate the location of the passage shock waves 
over a range of Mach numbers and operating conditions. Also, they allow qualitative 
comparisons of flow conditions tnrough blade passages for the two bladc shapes. How- 
ever, because oi the indeterminable effects of three -dimensional flow, data reduction 
techniques, and probe size on the accuracy of the static pressures,  a quantitative analysis 
of the data is not attempted. 
Flow 
INTRODUCTION 
Blade speed and blade loading of axial-flow compressor stages are continually being 
increased. in order to obtain a higher pressure ratio from incividual stages. As the local 
veiocities exceed the local sonic velocity, shock patterns o r  dr throughout the blade pas- 
sage. Associated with the shock formations are pressure losses in flow across  the shock 
as well as increases in viscous losses and other effects due to interactions with the blade- 
surface boundary layers. The magnitude of both these elTfectZ; increases with local flow 
Mach number and must be considered in an accurate appraisal o r  prediction of flow con- 
ditions across a blade row. 
The investigations reported in references 1 and 2 used both analytical and experi- 
mental approaches ' 7 identify the shock patterns that OCCUI- in high-speed rot?-ting corn - 
pressor blade rows. From these studies flow models for ;,redicting loss due to shock 
patterns ir, '3 blade passages were suggested. 
Recently. the Cewra l  Electric Conipany, a:der contract NAS3-7617 with NASA. 
experimentally evaluated the perforniaice of a number of high Mach number coinpressor 
rotors. The results of this investigation are reported in references 3 to 6. Suppleniental 
inst rumentation used during this investigation included high-frequeixy static-pressure 
transducers mounted in the outer casing above the rotor blade tip. From these nieasui-e- 
nients the static-pressure field in  the blad2 tip flow passages were constructed. 
In this report the static pressure contours in the blade passage at the tip of three 
rotor blade rows are presented for speeds between 90 and 110 percent of design and for a 
range of flow rates. The test  blade rows consisted of both double circular 31-c and niul- 
tiple circular arc blade shapes and were tested over a Mach nuiriber range of 1 . 2 3  to  
1. 59. A detailed analysis of the results is not attempted in this report. 
DESCRIPTION OF TEST ROTORS 
Detailed aerodynamic and mechanical design features of the three rotors used in this 
investigation are presented in reference 3. In this report, only significant blade design 
variations will be given. The three rotors represent two bkde  shapes and two levels of 
blade loading (indicated herein by D-factor) - The blade loading pai*anieter D-factor is 
developed in reference 7. A value of tip D-factor of 0. 35 is considered to be a medium 
level of loading. The rotor designated as 11-D had blade sections in the t ip region that 
were the familiar double circular arc form and had a tip D-factor of 0.45. The blade sec- 
tions in the tip region of the other two rotors used a inuitiple circuiar a r c  blade shape. 
This type of blade section perinits one rate of turning along the forward portion of the 
blade and another along the r ea r  portior and thus permits some control of blade surface 
Mach number. Rotor 11-B had the same D-factor as rotor II-D (D = 0.45), bL t with 
less  turning in the forward part of the blade. The rotor I-B had an inlet portion the same 
as B-B but the design D-factor was lower. Table I summarizes the principal blade design 
variations and blade speeds at which data are presented. 
ACQUISITION AND REDUCTlaN O F  DATA 
Local pressures  over the rotating blade tip were obtained by combining (1) measure- 
ments of average static pressures  from outer wall static-pressure taps with (2) nieasure- 
ments of the change in static pressure as obtained from high-response pressure t rans-  
ducers. These measurements were obtained from 10 static-pressure taps (covering the 
axial extent of the blade tip section) and 10 high-frequency pressure transducers (cover- 
ing the same axial extent as the wall taps), which were mounted in the outer wall casing 
(see fig. I). The transducers were of the quartz crystal type with a pressure sensitive 
2 
surface of approximately 0 .22  inch (0. 56 cni) dianieter. The t r~ i sducc r - r eco rd ing  sys- 
tem had a frequency response of over 100 000 hertz. The changes in  static pressure,  as 
sensed by the crystal transducers, were displayed on an oscilloscope. Using an elec- 
tronic triggering system, activated by ;i particular blade, a t race covering the static- 
pressure variation across  two blade passages is displayed. The triggering device also 
insures that the same two blaGe passages are sampled each time. From the known blade 
tip section geometry and blade speed, the pressure trace and blade position can be re-  
lated. 
A trace of a typical pressure change across  two adjacent klade passages is shown in 
figure 2. The blips over the trace are signals from the triggering device and were used 
to locate the trace wltli respect to the blades. On the trace,  sharp pressure changes 
measured across  the blade tip section (pressure to  suction surfaces) and across  a pas- 
sage shock allow an approximate positioning of the shock in the flow passage. 
Local values of static pressure on a given t race were obtained by determining an 
average value for each t race and relating it to the static pressure measured by the outer 
wall taps. ?->e local pressure values were positioned in the blade passage by plotting the 
trace of absolute pressure with passage width as shown in figure 3. When this process 
was completed for the 10 crystal transducers used, points of equal pressure were con- 
nected to form pressure isobars. An example of the pressure isobar plots is shown in 
figure 4. 
applied generally to  the static pressures  presented. The sample t race shown in figure 2 
illustrates several  of these factors. Firs t ,  the width of the band shown in figure 2, in- 
dicates the variations of pressure with time. Second, the t race also indicates that the 
prespure variations differ in the two adjacent blade passages. The data reduction pro- 
cedure averages these time and space variations. Average values vary from the extreme 
values by as much as *12 psi (1.03 N/cm ). Another indeterminable effect cn the ac- 
curacy of the local pressure is the diameter of the pressure sensitive face of the probe. 
The pressure transducer used in this investigation had a diameter of 0 .22 inch (0.56 cm), 
o r  nearly 10 percent of the passage width. It is assuraed that the readout value is some 
average of the pressures over the entire a r e a  of probe face. In particular, this may re- 
sult in smoothing out pressure peaks. Additional problem a r e a s  include the accurate 
alinement of the blades with the pressure t races  and the effect of outer casing boundary- 
layer flows on ineasured pressures.  
A number of factors make it difficult to  estimate an accuracy value that could be 
1 2 
INTERPRETATION O F  DATA 
The pressure isobar plots (such as shown in fig. 4) a r e  presented as essentially 
3 
two-dimensional maps. It should be recognized that there are three-dimensional flows 
present that have some indeterniinable effect on the static pressure patterns measured. 
Sources of these three-dimensional flows inciude Cur-Tature in th2 casing wall, boundary- 
layers on the casing and blade surfaces,  and blade tip clearance flows. 
The positior .i' shock waves in the blade passages is evidenced by sharp increases 
in static pressure and indicated by closely spaced pressure isobars. In particular, a 
bow wave lying just upstream of the blade leading edge and a passage shock can usually 
be identified. 
A measure of the sti-ength of the shock wave is indicated by the static pressure rise 
across  the shock. Shock strength is related tc the local Mach number of the flow and the 
angle between the flow direction and shock wave. 
ROTOR PERFOIZMANCE 
Design blade tip speed for  all rotors was 1400 feet per  second ( 6 .7  ni/sec), ant 
inlet relative Mach numbers at the blade tip was near 1.4. Thc overall performance of 
the three rotors is presented in figure 5. Mass-averaged valuzs of pressure ratio and 
efficiency a r e  plotted as functions of corrected weight flow. The readiqg number printed 
beside each data point affords a quick reference to the high-frequency data to be pre-  
sented and to blade element performance data in references 4 to 6. The effects of blade 
loading, blade speed, and blade shape on overall performance are all evident in the fig- 
ures  and are discussed in detail in references 4 to 6. 
PRESENTATION OF STATIC PRESSURE PLOTS 
Nineteen maps of the static pressure fields over the tips of rotating blade rows a r e  
presented in figures 6 t o  10. Each figure displays pressure contours for a given speed 
and over a range of flows o r  back pressure.  Par ts  (a) of these figures represent the 
flow conditions at the flow rate  nearest t o  stall, and parts (d) (part (c) of fig. 10) show 
flow conditions at the lowest back pressure measured (near choke). The numbers on 
each contour is the static pressure in psia. Thus, figure 6 shows pressure maps of four 
flow conditions in the 2D rotor at 100 percent speed, figures 7 and 8 are for the 2B and 
1B rotors,  respectively, a t  1CO percent speed, and figures 9 and 10 a r e  for  the 1B rotor 
at speeds of 110 and 90 percent of design. 
Table lI l ists  pertinent information for the various pressure maps. This includes 
the inlet relative Mach number of the flow entering the blades, the suction surface in- 
cidence angle, the inlet static and relative total pressures,  the outlet relative Mach nuin- 
4 
ber ,  and the outlet static pressure.  These were obtained by extrapolsting the data in the 
references to the tip radius or  by extrapolating the data from static pressure taps in the 
outer casing. Also included a r e  the reading numbers corresponding to  those on the per-  
formance curves of figure 5. 
blades) and rotor 2D (double circular arc blades) provides an opportunity to qualitatively 
evaluate the effect of blade shape on blade suction surface Xach number. One purpose of 
the reduced camber in the blade leading edge region, as designeci into the niultiple eir- 
cular arc blade shape, was to reduce the Mach number of the flow on the suction surface 
upstream of the shock location. Both blade sections operated at approximately the same 
level of inlet Mach number and had the same overall desibm blade loading. The generally 
lower levels of minimum static pressure measured for rotor 2D (fig. 6) as compared 
with those for rotor 2B (fig. 7) indicate that higher suction surface Mach numbers did 
occur with the double circular arc blade shape. In this design the multiple circular a r c  
blade shape did reduce the suction surface Mach number upstream of the shock. Coni - 
parative relative Mach number can lie computed from the minirnuin static pressures  and 
the blade suction inlet relative total pressures.  
A comparisnn of the minimum static pi essures  for rotor 2B (multiple circular arc 
SUMMARY 133MARKS 
The static pressure fields in the blade tip region of these high-speed rotor blade 
rows ) p r a t i n g  over a range of Mach numbers and flow rates are presei2ted in the form 
of static pressure contour plots. Inlet relative Mach numbers varied from 1.23 to 1. 59. 
Flow rate varied from the choke value to  that near blade stall. Results from both mul- 
tiple circular arc and double circular arc blade shapes ;re presented. 
Because of the indeterminable effects of three-dimensional flows, annulus boundary 
layers,  data reduction techniques, and probe size on the accuracy of the reduced static 
pressures,  a quantitative evaluation of this single set  of data is not attempted. However, 
the results do provide additional insight into the transonic flow conditions across  high- 
speed rotor blade rows and some qualitative and comparative conclusions can be reached. 
The reaction of the passage shock to the blade row back pressure,  the significant inter-  
action of shock with blade surface boundary layer,  and the ability of the multiple circular 
a r c  blade to minimize the shock '?let Mach number are indicated. 
Lewis Research Center, 
National Aeronautics and Space Administration, 
Cleveland, Ohio, October 5, 1970, 
720-03. 
5 
REFVRENCES 
Blade speed. 
designation percent of design I Rotor I 
1. Scliwenk, Francis C. ; Lewis, George W. ; and Hartmann, Melvin J. : A Preliminary 
Analysis of the Magnitude of Shock L.osses in Transonic Compressors. NACA RM 
E57A30, 1957. 
2. Miller, Genevieve R. ; Lewis, George W. , Jr.  ; and Hartmann, Melvin J. : Shock 
Losses in  Transonic C m p r e s s o r  Blade Rows. Paper 60-WA-77, ASME, Nov. 1960. 
3. Seyler, D. R. ; and Smith, Leroy H . ,  Jr. : Single Stage Experimental Evaluation of 
High Mach Number Compressor Rotor Blading. Part 1: Design of Rotor Blading. 
Rep. R66FPD321, pt. 1, Geneial Electric Co. (NASA CR-54581), Apr. 1, 1967. 
4. Seyler, D. R. ; and Gostelow, J. P. : Single Stage Experimental Evaluatjon of High 
Mach Number Compressor Rotor Blading. Par t  2: Performance of Rotor 1B. Rep. 
R67FPD236, General Electric Co. (NASA CR-54582), Sept. 22, 1967. 
5. Krabacher, K. W. ; and Gostelow, J. P. : Single Stage Experimental Evaluation of 
High Mach Number Compressor Rotor Blading. Part 4: Performance of Rotor 2D. 
Rep. R67FPD276, General Electric Co. (NASA CR-54584), Oct. 6 ,  1967. 
6. Gostelow, J. P. ; and Krabacher, K. W. : Single Stage Experimental Evaluation of 
High Mach Number Compressor Rotor Blading. Part 5: Performance of Rotor 2B. 
Rep. R67FPD278, General Electric Co. (NASA CR-54585), Oct. 13, 1967. 
7. Lieblein, Seymour; Schwenk, Francis C. : and Broderick, Robert L,. : Diffusion Fac- 
tor  for  Estimating Losses  and Limiting Blade Loadings in Axial-Flow -Compressor 
Blade Elements. NACA RM E53D01, 1953. 
DesiLm blade Blade 
loading tip sha!~es 
TABLE I. - PRINCIPAL BLADE DESIGN VARIATIONS 
I-B 
11-B 
I1 -D 
110 0.35 M CA 
100 
90 
100 0 . 4 5  M CA 
100 0 .45  DCA 
6 
TABLE 11. - FLOW CONDITIONS FOR PRESSURE CONTOUR MAPS 
1 . 3  
. 6  
. 4  
. 4  
12.63 
12.31 
12.07 
11. 30 
29.10 
29.10 
28.74 
26.91 
0.756 
.802 
.865 
1.248 
3. 8 
1. 8 
1.1 
. 9  
2 . 4  
1 .8  
1.2 
. 8  
13.12 
12.64 
12.14 
11.79 
13. 12 
12.80 
11.89 
11.28 
.- 
3 . 9  
1 . 4  
. 5  
~~ 
13.58 
12.96 
12.49 
Outlet stati: 
p re s su re .  
p2 F 
psi 
20.24 
19.97 
18.15 
13.20 
- N/cm2 
8.62 
8 .43  
7 .93  
7. 52 
2 K / m n  
13.95 
13. 76 
12. 51 
9 .09  
psi 
42.98 
42.11 
39. 83 
38.28 
1.454 0 . 1  12 .51  
1.455 -. 3 12.24 
1.453 -. 8 11.51 
1.468 -. 9 10.92 
29.61 0.736 
29 .01  
27.44 
26.37 1. 302 
8. 70 
8 .48  
8. 32 
7. 79 
42.23 
42.23 
41.71 
39.0'; 
21.03 
20.30 
19.42 
13 .33  
14.49 
13.99 
13.38 
9. 18 
1.435 
1.453 
1.458 
1.458 
1.384 
1.413 
1.424 
1.426 
9.04 
8. 71 
8. 36 
8. 12 
40.82 
40.97 
39.96 
38.92 
28. 12 0.  756 
28 .23  , 1 .887 
27.53 i 1.012 
26.82 I 1 . 2 3 3  
21.11 
20.10 
18. 19 
73.95 
22.54 
21.76 
20.35 
13.89 
19.43 
17.87 
14.00 
-
-
- 
14. 54 
13.85 
12.53 
9 .61  
15.53 
14.9Y 
14.02 
9.57 
13.39 
12.31 
9 .65  
-- 
1.558 
1.565 
1. 550 
1.587 
0.792 
.840 
.950 
1. 348 
9.04 
8 .  82 
8. 19 
7.77 
9. 36 
8. 93 
cl 6 1  
52.42 
51.72 
49.07 
47.02 
34.33 
34.48 
33.68 
36.12 
35.64 
33.81 
32.40 
23.76 
0.  704 
,879 
1. 114 
. . ._ . -- 
1.232 
1.270 
1.280 

Passage width 
Figure 3. . Positioning of absolute pressures in blade passage. 
1? 
12 / 
Figure4. - 
',2 1,2 
Example 3f static pressure coqtours across blade passage. 
9 
Equivalent v,eight flow, lbfsec 
(a) Rotor 18. (b) Rotor 28. (c) Rotor 20. 
Figure 5. - Overall performance of test rotors (from refs. 4. 5, and 6) .  
10 
pressure,  
psia 
c8 14' 
(a1 Reading 9; n e a r  stall. 
(c) Reading 14. 
12 
12 
(d) Reading 11; n e a r  choke. 
F igu re  6. - Pressu re  con tou rs  fo r  ro to r  20 at 1OC percent  of 
des ig  t i  speed. 
11 
Static 
pressure, 
16 
(a) Reading 53; near stall. 
(b) Reading 19. 
15 
( c )  Reading 18. 
(d) Reading 14; near choke. 
Figure 7. - Pressure contours for rotor 26 at 100 percent of 
design speed. 
12 
15 Static 
17 
II 12 
12 
{ r l )  Reading 65; near choke. 
Figure 8. - Pressure contours lor rotor 1B of 1012 percent 
of design speed. 
13  
(a) Reading 102 near stall. 
(b) Reading 101. 
12 12 ,12 
(c! Reading 95. 
(d) Reading 96; near choke. 
Figure 9. - Pressure contours of rotor 1B at 110 percent of design speed. 
14 
13 
ta) Reading 73; near  stall. 16 
i b) Reading 67. 
(c) Reading 69; near  choke. 
F igure 10. - pressbye c o n t o u r  for rotor 1 B  at 90 percent  of design speed. 
NASA--ley, 1971 - 1 E -5875 15 


Static-pressure contours in the blade passage at the tip of several high Mach number rotors

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