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RESEARCH MEMORANDUM _ for the

Bureau of Ordnance, Department of the Navy

EFFECT OF FIRST-STAGE BLADE DESIGN ON PERFORMANCE

OF MARX.25 TORPEDO POWER PLANT ,’ L./

By Harot% J. Schum and Jack W. Hoyt

Lewis Flight Propulsion Laboratory ’ Cleveland, Ohio

n:

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON

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Restriction/Classification Cancelled



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RESEARCH MEMORANDUM _ for the



OF MARK 25 TORPEDO POWER PLANT V ---../

By Harold J. Schum and Jack W. Hoyt

Lewis Flight Propulsion Laboratory , Cleveland, Ohio

CLASSIFIED DOCUMENT

This document contains c1assi!1ed information affecting the National Defense of the United States within the meaning of the Espionage Act, USC or the revell r to an unauH ry law.

Infor Iparted only I I naval servtc lprlate civill~ 'ederal Gover lterest . there! Imown loyaltjr ""'" .... L""'"~"' .. ~V .... "V ............... ~ ... J' .. lust be informed thereof.

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON

RES'

NACA EM SE9G20 1111111111111111~~~I[~flljmll[~[~1111111111111111 3 1176014379490

NATIONAL ADVISORY COMMITl'EE FOR AERONAUTICS

RESEARCH MEMORANDUM

for the



OF MARK 25 TORPEDO POWER PLANT

By Harold J. Schum and Jack W. Hoyt

SUMMARY

The effect of rotor-blade length, inlet angle, and shrouding was investigated with four different nozzles in a single-stage modification of the Mark 25 aerial-torpedo pOW'er plant. The results obtained with the five special rotor configurations are compared with those of the standard first-stage rotor with each nozzle. Each nozzle-rotor combination was operated at nominal pressure ratios of 8, 15 (design), and 20 over a range of speeds from 6000 rpm to the design speed of 18,000 rpm. Inlet temperature and pressure conditions of 1000° F and 95 pounds per square inch gage, respectively, were maintained constant for all runs.

The increase in annular flOW' area, either through greater blade length or wider inlet angle, apparently resulted in the greatest increase in efficiency. The rotor having longer blades showed slightly greater efficiency than the blades with wider inlet angles with each nozzle at the high pressure ratios. Removal of the blade shroud caused a decrease in efficiency as expected. Only part of this decrease in efficiency was recovered when a close-fitting stationary shroud cap was placed over the active part of the blading.

The lI8Ximum brake efficiency of 0.54 was obtained when the longest blades (0.45 in.) were used with a cast sharp-edged-inlet nozzle (H). When this configuration was run as a two-stage unit with a standard second-stage rotor, hOW'ever, the resulting efficiency was less than that observed when both standard rotors were used. This decrease in efficiency was probably caused by less favorable outlet velocities from the 0.45-inch first-stage rotor due to the increase in flow area, thereby reducing the effectiveness of the second stage.

R

NACA EM SE9G20 2

A water-channel investigation of X5-scale models of the standard 170 -inlet-angle blade and the special 200 -inlet-angle blade indicated the presence of a normal shock at the blade inlet, causing

'i' both subsonic flow through the blade passages and a loss in total f' !.~ pressure. ~~ 'I',

INTRODUCTION

At tlie request of the Bureau of Ordnance, Department of the Navy, the performance of a turbine from the Mark 25 aerial-torpedo power plant is being investigated at the NACA Lewis laboratory. Although the gas turbine is composed of two counterrotating stages, the second rotor is used primarily for balance. For possible application of high-pressure single-stage turbines to rocket accessory drive, a Mark 25 turbine with the second-stage blading removed was investigated. The over-all performance of the standard twostage turbine and the first-stage turbine with five nozzles is reported in references 1 and 2 , respectively, and a study of axial nozzle-rotor clearance in reference 3.

Because of the geometry of the turbine installation in the Mark 25 torpedo, no precise survey measurements could be made near the nozzles and the blades during these previous investigations. Only the over-all performance of various combinations of nozzles and blades could be used to obtain information leading to performance improvements.

The investigation reported herein was undertaken to obtain information regarding the variation of the over-all efficiency of turbines of this type with nozzles designed to vary the weightflow rate and blades designed to change the effective flow area through the passages.

Four different turbine nozzles were investigated in conjunction with modifications to the first-stage blade design. These modifications included. special blade heights of 0.35 and 0.45 inch as compared with the standard blade height of 0.40 inch. A third special first-stage rotor was fabricated with a blade-inlet angle of 200 instead of the standard 170 , the blade height being the same as that of the standard blade. When the results obtained 'with these three rotors are compared with the results previously obtained with the standard rotor, the effect of varying the effective flow area through the blading can be determined. The use of unabrouded blading facilitates turbine manufacture and hence the extent of the change in over-all performance was ascertained with the shroud

NACA SE9G20

removed from the 0.35-inch rotor blading. Turbine performance was also determined with a close-fitting stationary shroud cap installed around the active portion of the blading to reduce the air leakage around the blade tips • Finally, the most efficient combination of nozzle and first-stage rotor was investigated in a two-stage Mark 25 turbine with use of the standard second-stage rotor.

All nozzle-rotor configurations were investigated at pressure ratios of 8, 15 (design), and 20 and a range of turbine speeds from approXimately 6000 to 18,000 rpm at inlet temperature and pressure conditions of 10000 F and 95 pounds per square inch gage, respectively. In order to evaluate the actual output of the nozzle-rotor combinations, the windage losses of each blade design w,ere determined.

Water-channel investigations applying the hydrauliC analogy to supersonic air flow in order to study the flow through the firststage blades at simulated operating velocities were made of both the 200 - and the standard l70 -inlet-angle blades.

TURBINE MODIFICATIONS AND SETUP

The standard Mark 25 torpedo power plant is a two-stage counterrotating turbine with partial gas admission. The two stages operate at the same rotative speed but in opposite directions in order to eliminate gyroscopic action. Design speed is 18,000 rpm. A description of the turbine and the modii'ications necessary to convert this unit into a single-stage turbine are described in references 1 and. 2.

Blades. - The standard first-stage turbine bas an outer diameter of 11.00 inches and 0.40-inch shrouded blades with an inlet angle of 170 • The blades of a sta.nda.rd first-stage rotor were machined to 0.35 inch and. a steel shroud was shrunk around the wheel to make one speCial turbine rotor. The second special rotor had blades 0.45 inch high and was designed so that the pitch diameter was not changed from that of the standard rotor; that is, the tip diameter was increased and. the hub diameter was decreased 0.025 inch. The third special rotor used had a 200 inlet angle and the standard blade height of 0.40 inch. .

Shrouds. - The effect of shrouding was determined by removing the steel shroud band around the 0.35-inch blades. Because unshrouded turbine wheels are more easily fabricated, determination of the aerodynamic penalty involved in the use of such wheels was considered

~ , r-'

NACA RM SE9G20 4

desirable. A close-fitting stationary shroud cap was installed around the active part of the unshrouded 0.35-inch blades. The stationary shroud cap is shown with the unshrouded rotor in figure 1.

Nozzles. - The alphabetical notation of the nozzles used in the investigation was arbitrarily selected to distinguish nozzle design. Nozzles A, E, and H are discussed in references 1 and 2. Nozzle A has rounded inlets to the rectangular converging-diverging nozzle ports. Nozzle E has reamed ports with rounded inlets. These ports are cylindrical with no divergence in the nozzle, completed gas expansion occurring in the clearance space between the nozzle and the blades. Nozzle H has rectangular ports with sharp-edge inlets to the throats cast by a dif'ferent technique from that used for nozzle A. Nozzle I (not previously investigated) is similar to nozzle H in design but has smaller inlets and greater wall divergence in the ports. Both nozzles H and I utilize the gas-jet contraction effect caused by the sharp inlet edges, producing the equivalent of converging-diverging flow. The throat of the two nozzles is thus at the vena contracta, from which the gas expands to the nozzle outlet. All four nozzles have nine ports with 90o-arc admission. Additional characteristics of the nozzles are given in the following table:

.------- ---~----Nozzle Total Observed

measured air throat weight area flow

(sq in.) (lb/hr) --- -~--~~---

A 0.183 955 E .193 972 H .226 1132 I .217 995

-~-----~-

The nozzle-throat area and the air-weight flow are not proportional, probably because of variance in vena contracta with d1:f'ferent nozzle designs.

The axial nozzle-rotor clearances were set to give a running clearance of 0.030 inch at operating temperature. Radial nozzle settings as well as the proper allowances for thermal changes in axial clearances were made to con:f'orm with the recommendations of reference 1.

NACA EM SE9G20 5

In order to determine what gains in a two-stage Mark 25 torpedo turbine could be realized by using the most efficient single-stage nozzle-rotor coni'iguration, a turbine was assembled to operate with a standard second-stage rotor in combination with the 0.45-inch blading and nozzle H.

The shock pattern and the possible flow separation through the standard 170 -inlet-angle blades and the special 200 -inlet-angle blades were studied by placing X5-scale models of the two pitch-line sections in a water channel similar to that described in reference 4. The water channel essentially consists of a trough of shallow water flowing at the velocity necessary to simulate a given air-flow Mach number. The sur:race waves caused by the blades in the water channel represent shock waves in air. The waves are photographed by illuminating the channel from underneath and focusing the shadows of the waves on a ground glass. The theory of the hydraulic analogy to compressible air flow is given in references 4 and 5.

PROCEDURE

For each nozzle-rotor combination, the effect of pressure ratio and blade-jet speed ratio on turbine efficiency was investigated. These investigations were made at pressure ratios of 8, 15 (design), and 20 over a speed range from 6000 to 18,000 rpm with a constant inlet temperature of 10000 F and a constant 'inlet pressure of 95 pounds per square inch gage. Because of flow limitations of the outlet ducting, the maximum pressure ratio obtained with nozzle H was 19.

The rotation losses of the single-stage turbine with the standard rotor are presented in references 1 and 2 and include an individual evaluation of the mechanical friction losses, the disk windage losses, and the losses incurred by the air-pumping effect with partial admission. Rotation losses of the 0.35-inch unshrouded blade were determined by measuring the power required to motor the rotor at speeds from 6000 to 18,000 rpm and at various air densities in the turbine case. Because of stress limitations acquired when the outer shroud was shrunk onto the blades, motoring the shrouded 0.35-inch blades under cold-air conditions for which there would be no compen~ting thermal effects was considered u.nsa:f'e. The windage losses for this rotor were therefore calculated from previous data.

NACA EM SE9G20 6

CALCULATIONS

The brake, the rotor, and the blade efficiencies were computvd according to the method given in reference 1. Brake efficiency is the ratio of the brake power calculated from the torque and the speed at the dynamometer shaft to the available isentropic power based on inlet total temperature and pressure and outlet static pressure; rotor efficiency is the ratio of the brake power plus the mechanical losses in the gears and the bearings to the isentropic power; and blade efficiency is the ratio of the brake power plus the mechanical losses and the disk and blade windage losses to the isentropic power. Blade-jet speed ratio is defined as the ratio of the blade speed at the pitch diameter of the rotor to the ideal nozzle-jet velocity based on isentropic expansion from the inlet total temperature and pressure to the outlet static pressure. Pressure ratio is defined as the ratio of the inlet total pressure to the outlet static pressure.

Charts of the windage and mechanical losses of the single-stage unit prepared from motoring studies of a disk (reference 2) and the complete rotor are based on the analysis and the calculation procedure described in reference 1 for the two-stage turbine. The power required to motor the unshrouded O.35-inch blades at various turbine speeds and temperatures and pressures in the turbine case is given in table I. Because of stress limitations, the power necessary to motor the O.35-inch shrouded blades was necessarily computed from data for the O.40-inch-blade-height rotor (refer-ence 2) by assuming that the blade windage loss was proportional to 11 •5, where 1 is the length of the turbine blade (reference 6). Data obtained with the shrouded O.35-inch blades were calculated both with and without this correction and the agreement with the basic data obtained with the O.40-inch standard rotor was found to be within experimental accuracy. On this basis and because the pitch diameters and the outside diameters of the rotors with the O.45-iDch and the O.40-inch blades described in reference 2 are almost the same, the two rotors were assumed to have the same windage losses. Because the nozzles had 900 -nozzle-arc gas admiSSion, one-fourth of the rotor blades are active and hence not subject to windage or pumping loss. Because windage loss of the unit is the total disk windage (reference 2) plus three-fourths of th-a windage due to the inactive rotor blades, figure 2 was prepared with the additional loss due to the inactive O.35-inch shrouded and unshrouded blades considered. This figure may be used. directly to find the windage and mechanical losses in the O.35-inch-blade turbine if the gas density in the turbine case is known. A silnilar 'chart for the rotor with O.40-iDch blades is presented in reference 2.

NACA EM SE9G20 7

For the water-channel studies of the blade shapes, the inlet Mach number to the blade under turbine operating conditions was calculated for the design pressure ratio of lS through the nozzle, assuming a nozzle velocity coefficient of 0.96, an inlet temperature of 10000 F, and design turbine speed of 18,000 rpm. This Mach number and the design inlet angle of flow were simulated in the wa t.er channel.

RESULTS AND DISCUSSION

Single-Stage Performance with Various

Nozzle-Rotor Combinations

Efficiency data for the rotors having 0.3S-inch unshrouded blades, 0.3S-inch unshrouded blades with a stationary shroud cap, 0.3S-inch shrouded blades, standard 0.40-inch blades and nozzle I, 0.4S-inch shrouded blades, and for the runs with the 200-inlet blades are presented in tables II to VII.

Nozzle A. - The brake efficiency of the single-stage turbine with nozzle A and the various blade designs is shown in figure 3. The data of reference 2 for the standard 0.40-inch rotor are also shown for comparison. At a pressure ratio of 8, the standard rotor is the most efficient for this nozzle. When the pressure ratio is increased to the design value of lS or to 20, the two rotors having a greater flow area (the 0.4S-in. blade and the 200-inlet blade) showed improved performance. Figure 3 also shows the reduction in performance when the blade height is reduced to 0.3S inch. 'l.'he very substantial drop in efficiency when the shroud band is removed is only partly recovered when a close-fitting stationary shroud cap is placed over the active portion of the blading. Because the difference in windage losses in the shrouded and unshrouded 0.3S-inch blades is small (fig. 2), the drop in efficiency of approximately 0.04 caused by the two rotor configurations at maximum blade-jet speed ratio and the design pressure ratio of lS indicates high tip losses when the unshrouded wheel is used. The corresponding drop in efficiency at pressure ratios of 8 and 20 were approximately 0.08 and O.OS, respectively.

Nozzle E. - When nozzle E is used (fig. 4), there is little difference in performance between the standard 0.40-inch rotor, the 0.4S-1nch rotor, and the 200-inlet-angle rotor, especially at the higher pressure ratios. The performance with the 0.3S-inch rotor shows the same trends as those for nozzle A.

NACA EM SE9G20 a

Nozzle H. - A gain in performance of approximately 0.05 over the standard 0.40-inch blade was obtained with nozzle H and. the 0.45-inch blade (fig. 5). The maximum obtained brake efficiency of 0.54 occurred at a blade-jet speed ratio of 0.295 and a pressure ratio of a. The peak brake efficiency, which would be at a higher blade-jet speed ratio, could not be determined because of the la,OOO-rpm speed limitation of the turbine. The same limitation of peak efficiency occurs with all the nozzle-rotor configurations investigated. The 200 -inlet-angle rotor showed substantially the same performance as the standard rotor with nozzle R. This nozzle handled the largest air flow of the nozzles operated, indicating that a greater flow area through the turbine blade is advantageous for more extraction of power from the single-stage turbine.

Nozzle I. - The results of the various rotor configurations with nozzle I are presented in figure 6. At the higher pressure ratiOS, the 0.45-inch blade shows the most efficient performance. The unshrouded 0.55-inch blades again had the lowest performance.

In general, the rotor and blade efficiencies presented herein for the various nozzle-rotor configurations follow the same trends as the brake efficiencies.

Two-Stage Turbine with Most Efficient

Nozzle-Rotor Combination

In general, experience with the various types of nozzle and blade in the single-stage turbine indicates that a som~hat larger flow area in the rotor is advantageous for the extraction of greater power from the first stage. Results have shown that the rotor having longer blades than that of the standard rotor had slightly greater efficiency than the blades with increased inlet angles with each nozzle. In the event that the flow leaving the rotor is subsonic, however, the greater area in the first stage may cause the velocity leaving this stage to be insufficient for developing as much useful work from the second stage as that observed during the investisation of the standard Mark 25 two-stage turbine. Study of the standard first-stage turbine blades indicated that the throat area was insufficient to swallow a supersonic air stream under operating conditions. Subsonic flow appeared likely to prevail at the blading outlet. Velocity diagrams based on this premise indicated both low outlet velocity from the first stage and unfavorable inlet angle for the second. stage. In order to investigate the possible adverse effects of increasing the flow area of the firststage rotor, the most efficient first-stage combination (O.45-in.blade rotor and nozzle H) was therefore operated with the second

NACA EM SE9G20 9

stage. It should be noted, however, that at a blade-jet speed ratio o:f 0.20, which would be approximately the peak performance point for the two-stage turbine, the 0.45-inch-blade rotor gives only a slight gain in performance over the standard rotor.

In figure 7, the performance of the two-stage turbine with the C.45-inch-blade first-stage rotor and nozzle H is compared with the sGandard Mark 25 blading with nozzle H. The lower efficiency of the unit with the higher blade-height first stage may be attributed to unfavorable inlet conditions of the second stage. Because the windage losses for the 0.40- and 0.45-inch blade-height first-stage rotors are considered the same, the windage losses for the twostage units using these rotors are the same. The blade efficiencies therefore bear the same relation to each other as do the brake efficiencies for the two-stage units. The gain in first-stage output is apparently more than outweighed by the drop in second-stage performance. The data for the two-stage unit with the 0.45-inch bladeheight first stage is presented in table VIII.

Water-Channel Studies

In order to ascertain if' supersonic velocities are maintained throU8h the turbine-blade passages, XS-soale models of the standard 170- and special 200-inlet blades were placed in a water channel to investigate the flow by the hydraulic analogy. Photographs of the flow throU8h models o:f the turbine blades taken in the water channel at an analogous Mach number o:f approximately 1.6, corresponding to assumed turbine operating conditions at a speed of 18,000 rpm and an inlet temperature of 10000 F, are shown in f igure 8. A nozzle pressure ratio of 15 and a nozzle-velocity coefficient of 0.96 were assumed for the Mach number calculations. The leading edge o:f the 17o-inlet-angle blades was placed at an angle o:f attack of 00 and the 200-inlet-angle blades at an angle o:f attack o:f 30 in the water stream.

In each set of blades, a wave corresponding to a normal shock in air appears at the inlet to the passage between the blades. The wave appears to be somewhat farther inside the passage in the 200-inlet-angle design. The occurrence of a normal shock in the inlet of the blade passage is indicative that the throat area is insufficient and that a rise in pressure together with substantial blade losses would result. That the analogous velocity inside the passage is subsoniC is evidenced by the apparent !nabili ty of the blading to swallow the shock and by the lack of oblique waves inside

NACA EM SE9G20

the passage. For maximum blade performance in supersonic turbine passages, it would be more advantageous to allow the normal shock to be swallowed completely. This experiment also supports the conjecture that subsonic flow prevailed at the rotor outlet.

SUMMARY OF BESULTS

10

The performance of the first stage of a Mark 25 two-stage turbine with several blade designs was investigated with each of four nozzles. The most efficient first-stage blade and nozzle combination was operated with a standard second-stage rotor as a two-stage turbine. The first-stage blade designs were examined in a water channel for shock-wave formation. The following results were obtained:

1. A somewhat larger flow area in the first-stage blades than that provided in the standard Mark 25 turbine ;was found to be advantageous for more efficient extraction of power fram the first stage. At the high pressure ratios, the rotor having longer blades than that of the standard rotor showed slightly greater efficiency than the blades with increased inlet angles with each nozzle.

2. The highest first-stage brake efficiency was obtained with a rotor having blades 0.45 inch high and a cast sharp-edged-inlet turbine nozzle (H). For a pressure ratio of 8 and blade-jet speed ratio of 0.295, the maximum brake efficiency of the combination was 0.54, a gain of 0.03 over the standard Mark 25 first stage under the same conditions.

3. Removing the shroud band from the outside of the turbine blades caused a substantial drop in efficiency. Only part of this decrease in efficiency was recovered when a close-fitting stationary shroud cap "was placed over the active portion of the blading.

4. When the most efficient first-stage rotor and nozzle combination was assembled with a standard second-stage rotor to form a twostage turbine, the resulting brake efficiency was less than when a standard Mark 25 first-stage rotor was used. This decrease was presumed to be caused by unfavorable outlet velocities from the first stage and inlet angles to the second stage due to the greater bladeheight first-stage wheel, causing the second stage to be of reduced effectiveness.

5. Water-channel investigation of X5 models of the standard 170 - and the special 200-inlet-angle Mark 25 first-stage blade

NACA EM SE9G20 11

designs indicated the presence of a normal shock at passage inlet caused by insufficient passage area. shock caused the flow in the blading to be subsonic in a loss in total pressure.

the bladeSuch a normal and resulted

Lewis Flight Propulsion Laboratory, National Advisory Committee for Aeronautics,

Cleveland, OhiO, August 4, 1948.

Approved:

rl

Robert O. Bullock, Aeronautical Research

Scientist.

Oscar W. Schey, Aeronautical Research

Scientist.

~~ Harold J. Schum,

Aeronautical Research Scientist.

Jack W. Hoyt, Aeronautical Research

Scientist.

NACA EM SE9G20 12

1. Hoyt, Jack W., and Kottas, HaITY: Investigation of Turbine of Mark 25 Torpedo Power Plant with Five Nozzle Designs. NACA EM SE7I03, Bur. Ord., 1947.

2. Hoyt.1 Jack W.: Investigation of Single-Stage Modified Turbine of Mark 25 Torpedo Power Plant. NACA EM SE7L15, Bur. Ord, 1948.

3. Hoyt, Jack W • .1 and Kottas, HaITY: Eff'ect of Turbine Axial Nozzle-Wheel Clearance on Performance of' l-Brk 25 Torpedo Power Plant. NACA EM SE8B04, Bur. Ord., 1948.

4. Johnson, R. H., Nial, W. R., and Witbeck, N. C.: Water Analogy to Two-Dimensional Air Flow. Rep. No. 55218, Res. Lab. and Gen. Eng. and Consulting Lab., Gen. Elec. Co. (Schenectady), Aug. 28, 1947. (U.S. Army Ord. Proj. Hermes (Tul-2000).)

5. Orlin, W. James, Lindner, Norman J., and Bitterly, Jack G.: Application of' the Analogy between Water Flow with a Free Surf'ace and Two-Dimensional Compressible Gas Flow. NACA Rep. 875, 1947.

6. Stodola, A.: Steam and Gas Turbines. Vol. I. McGraw-Hill Book Co., Inc., 1927, p. 200. (Reprinted, Peter Smith (New York), 1945. )

w --

NACA RM SE9G20

TABLE I - DATA FOR WINDAGE AND MECHANICAL

LOSSES OF SINGLE-srAGE TURBINE WITH

0.35-INCH UNSHROUDED BLADES

-Turbine Horsepower Air tem- Air press sure speed to motor perature in turbine

(rpm) turbine in tur- case bine case ( in. Hg abe.) (~)

6069 0.84 129 27·55 .82 132 22·59 .78 133 15.65 .68 132 9.36

I --- ~ -_.-

1.41 162 27.71 1.36 165 22.96 1.28 162 16.01 1.09 154 9.21

--

10,115 2.16 195 27·51 2.01 200 22.96 1.93 196 15·11 1.70 186 , 9.34

1

12'13

8 ~--- --._---

3.20 236 21.31 2.96 242 22.14 2.12 234 15.13 2.40 215 9.24

14,161 4.34 284 27.05 4.06 297 22.18 3.68 286 15.81 3.26 253 9.25

----~ ..

16,184 5.49 353 27.14 5.11 351 22.16 4.69 328 15.66 4.21 286 9.43

-- .------~----

18,201 1.32 381 26.81 6.90 388 22.44 6.30 351 15.11 5.52 320 9.44

13

TABL

E II

-EF

FICI

ENCY

DAT

A FO

R SI

NGLE

-STA

GE T

UBBI

NE W

ITH

0.35

-INC

H U

NSHR

OUDE

D BL

ADES

[In

let

te,m

per

atu

re,

1000

0 F

; in

let

pre

ssu

re,

95 I

b/sQ

. in

. g

age]

Noz

zle

Pre

s-A

ir

sure

w

eig

ht

rati

o f

low

Fue

l.

air

ra

tio

I Ho

rsep

wer

'fu

rbin

e ~1Ce

'Bla

de-

Gas

cI

en-

A

8

! A

15

A

. 20

E

8

(lb

/hr)

957.

4 I

956.

4 I

957.

4 95

8.4

: 95

8.4

: 95

7.4

I 95

7.4

I

952.

5 i

952.

5 95

2.5

952.

5 95

3.6

1

0.01

.37

.013

7 .0

137

.013

6 .0

136!

.0

137:

.0

137:

0.01

.37

.013

7 .0

137.

.0

137

.013

7, a

vail

ab

le

spee

d

ho

rse-

jet

sit

1 i

n

from

ise

n-

(rpm

) p

aver

ep

eed

turb

ine

• tr

op

ic

rati

o

case

ex

pan

sio

n

(lb

/cu

ft)

61

.26

6,07

9 13

.64.

0.

0985

0.

0312

61

.20

8,09

2 17

.07

.131

1 .0

317

61.2

6 10

,135

, 19.

87

.164

3!

.032

1 61

.30

12,1

78!2

2.15

: .1

974'

.0

327

61. 3

0 14

,161

230

75

.229

5 .0

330

61.2

6 16

,174

24.7

3 .2

621

.033

3 61

.26

18,1

8724

.87

.294

7 .0

335

1

73.5

6 73

.56

73.5

6 73

.56

73.6

5

6,09

9:14

.95

8,09

218.

72

10,1

25 . 2

1. 86

12

,128

i24.

42

I

14,1

41126

.38

0.09

00

.119

5 •

.149

5 .1

790

' .2

088

0.01

83

.018

6 .0

186

.018

7 .0

187

952.

5 .0

1.37

73

.56

952.

5 .0

137

73.5

6 16

,184

28.

11

I .2

389

.019

0 18

,237

29.

09

.269

2.

.019

2

953.

6 ,

953.

6 95

3.6

, 95

3.6

. 95

3.6

9530

6 95

3.6

0.01

.37

.013

7 .0

137

.013

7 .0

137

.013

7 .0

137

971.

4 ,0

.013

9 97

1.4

.013

9 97

0.7

.013

9 97

1.4

.013

9

78.7

3 78

.73

78.7

3 78

.73

78.7

3 78

.73

78.7

3

62.1

7 62

.17

62.1

3 62

.17

I

6,07

9 15

.03

8,10

2:18

.93

10,1

25'2

2.12

. 1

2,15

8 24

.76

i 14

,141

'26.

70

. 16,

194

28.1

8 18

,227

29.

25

0.08

68

.115

7 '

.144

6 .1

736

' .2

019

.231

3 .

.260

3

0.01

46

• 015

3 .0

155

.015

4 .0

157

.015

9 .0

162

I

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78.3

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78.3

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78.3

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C-2003411.20.47

dI

NACA RM SE9G20 37

i

i

Figure 1. - Front view of first-stage turbine of Mark 25 power plant with nozzle removedto show unshrouded 0.35-inch rotor blade and shroud cap.

I

NACA RM SE9G20 38

6 .--------,----,--- 1-- - - -r-- - - ,----- - -. ll~ur~-~n~:;;eed' rpm Blade

Shrouded Unsnrouded

-}---{---- -- -- - --- -- -. - /' 18,207

~Vi

-- 8,092

-

- I- -- - I- - J 6,069

I

~ o .01 .02 .03 .04 .05 .06 .07

Gas density, lb/cu ft

Figure 2. - Windage and mechanical losses of first-stage turbine with 90 0 _

arc admission and 0.35-inch shrouded and unshrouded rotor blades.

I

NACA RM SE9G20

>, <J c Gl

<J

'-Gl

(l)

-'" ~

m

.40

• 30 f----I'--~V

.20

.50 -

---I .40 I -- ---

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T

.40

.20 ~~---I---l-

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(a) Pressure rat i 0, S.

. -- -- -- -

-t, i

i ___ L_

. 1.

( b) Pressure rat io, 15.

height ( in.)

0 0.35 0 .35 A .35 "l .40

_I __ ~~ ~;~_ .16 .20 .24

Blade-jet speed ratio

(c) Pressure ratio, 20.

R ·TJ :---J1

Blade design

Unshrouded Shroud cap Shrouded Standard

l __

l

(reference 2) : Shrouded ~ 200 inlet L. _ 1 ._

.2S .32

Figure 3. - Effect of olade design on performance of singlestage turbine with nozzle A.

39

NACA RM SE9G20

'"' u

.60

.50

.40 .

.30

.20 -

• 50

~ .40 u

-'" ~ l" .30 co

.20

.10

. 50

.10 -.08 .12

o o 6.

-- V

<> +

.45

.40 I

(a) Pressure ratio, 8 .

(b) Pres5ure ratio, 15 •

.16 .20 .24 Blade-jet speed rat io

(e) Pressure ratio, 20.

Blade design

Unshrouded Sh roud cap Shrouded ~ Standard (reference 2) Shrouded 200 inlet

1_

r

I t ~,

I ~ I

J

... =a .28 .32

Figure 4. - Effect of Olade design on performance of singlestage turoine with nOlzle E.

40

'""-" I~

NACA RM SE9G20

>. 0 c: Ql

0

... Ql

~ ~

<D

.60

.50

.40

.30

.20

. 50

.40

.30

.20

.10

or .40·

(a) Pressure ratio, S .

I

Blade height ( in.)

0 0.35 I 0 .35

1 ~ ~t A .35 v .40

-j ;... ¢ .45 + .40

I

( b) Pressure rat i 0, 15.

Blade-jet speed ratio


I~n

£ Blade design

Unshroudcd Shroud cap Shrouded Standard ( reference 2) Shrouded fOo inlet

J ~-

--

-----

--0 ___ - ---

, --

.2S .32

Figure 5. - Effect of blade design on performance of singlestage turbine with nozzle H.

41

NACA RM SE9G20

>. o c: ., o

.60

.50

.40

.30 -

.20

. 50

.40

.30

.10

. 50

.40

.30

.20

.10 .08 .12

(a)

(b)

Pressure rat io, 8 .

Pressure rat io, 15 .

Blade design

Unshrouded Shroud cap Shrouded Standard

.45 Shrouded

.40 200 inlet

-~ -~

J

J

.20 J __ I _ _1-._--1

.16 .24 .28 .32 Blade-jet speed ratio


Figure 6. - Effect of blade design on performance of singlestage turbine with nozzle I.

42

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u c Q) u ~

~

Q)

Q)

:¥

«I

L ID

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1

6

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4

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ure

7

. -

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C-237767- 15-49

NACA RM SE9G20 44

(a) Standard blades with 170 inlet angle.

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(b) Special blades with 200 inlet angle.Figure 8. - Water-channel photographs of flow through pitch section of blades at analogous

Mach number of 1.6,

RESEARCH MEMORANDUM - Digital Library/67531/metadc64333/m2/1/high... · cations included. special...

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Transcript of RESEARCH MEMORANDUM - Digital Library/67531/metadc64333/m2/1/high... · cations included. special...