A set of modified NACA 65-series blower blades designed for axial inlet velocity, high inlet Mach number, and high blade loading was investigated for trailing-edge thickness of 0.015, 0.030, and 0.045 inch to determine the effect of trailing-edge thickness on single-stage axial-flow-compressor performance. Trailing-edge thickness effects were small except at the highest tip speed investigated (915 ft/sec). Trailing-edge thickness up to 30 percent of maximum blade thickness were used without sacrifice of performance of NACA 65-series blades

Moses, J J

Serovy, G K

English

NASA Technical Reports Server

RESEARCH MEMORANDUM 
SOME EFFECTS OF BLADE TRAILING-EDGE THICKNESS ON 
PERFORMANCE O F  A SINGLE-STAGE AXIAL-FLOW 
COMPRESSOR 
B y  J. J. Moses and G. K. Serovy 
'L 
Lewis Flight Propulsion Laboratory 
2LASSFF1CATiON CANCELLED 
Cleveland, Ohio 
"- "" """ ""_ """ "" 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
https://ntrs.nasa.gov/search.jsp?R=19930086736 2020-06-17T14:22:26+00:00Z
1 NACA RM E5lF28 
SOME EFFECTS OF BLADE TRAILING-- THICKNESS ON XRFORMNCE 
R N 
OF A SINGL?3-STAGE .KClXL-W COMpIiEsSOR 
By J. J. Moses and G. K. S e r a v y  
An Fnvestigation was conducted t o  determine some effects of blade 
trailing-edge thicknms on the perf OlZLYaZlce of a single-stage axial-fluw 
compressor haw a t i p  diameter of 14 inches and a hub-to-tip diameter 
ratio of 0.8 at the rotor leading edge. A r o t o r  row of modified NACA 
65-series blower blades designed for  high  inlet  Mach nuibers, high 
blade loading, and axial   in le t   veloci ty  was investigated for blade 
trailing-edge thicknesses of 0.045, 0.030 and 0.015 Inch, which gave 
ratios of trailing-edge  thickness.to msximum blade thickness of 0.30, 
0.20,and 0.I.D. Over-all total-pressure ratio and adiabatic efficiency 
t i p  speeds from 450 t o  915 feet per second. A range of re la t ive equiv- 
a lent   inlet  Mach number from 0.40 t o  0.93 at  the rotor mean rad ius  was 
covered; the  approx-iee Reynolds number based on blade chord varied 
from 261,000 t o  604,000. 
4 were determined for a c q l e t e  range of w e i g h t  flaws at six equivalent 
No measurable effect  of trailing-edge thickness on total-pressure 
ra t io  w&s obtained over the  efficient  operating speed range of the cam- 
pressor. The efficiency tended t o  increase slightly with increasing 
trailing-eQe thickness. The magnitude of the change, however, was 
within the accuracy of measurement (1 t o  2 percent). A lower c r i t i c a l  
Mach rider was indicated for the O.Ol5-inch thickness than for  the 
0.030- and 0.045-inch thicknesses. 
Trailing-edge thicknesses up t o  30 percent of maxFmum blade thick- 
ness were used without sacrifice of performance for the NACA 65-series 
blade sections. 
Axial-f low compressor-blade fabrication would be s-lified if the . permissible  thickness of the   blade  t ra iung edge could be  increased 
over the theoretically determined blade profile. IncreEts.ing the blade 
trailing-edge thickness m y  introduce a decrease in blade operating 
2 NACA RM E5lF28 
efficiency. The data of refereaces 1 t o  3 indicate that to obtain max- 
imum blade-section efficiency, the thfclmess of the t ra i l ing  edge 
should be as small as possible. On the other hand, no appreciable 
trailing-edge-thickness effect 011 blade performance for thicknesses up 
t o  20 percent of the maximum blade thickness is lndicated irn refer- 
ence 4. 
Results f o r  isolated airfoils presented i n  reference 5 show that 
an increase i n  l i f t  and an increase i n  profile  losses occur when 
trailing-edge thichness is increased. XP the lift Facreases, a 
decrease i n  blade-section lift-drag r a t io  all not  necessarily  result 
i n  a decrease in compressor eff'iciency because blade-profile  losses are 
only a small part  of the total  losses of a compressor. 'Ilherefore the 
investigation was conducted to determine the  effect of blade trail ing- 
edge thickness on the performance of a typical axial-flow cornpressor 
6 tage . 
A single row of compressor blades with modified NACA blower blade 
sections of 65-(14)Lo, 65-(9.6)09, and 65-(7.3)08 a t  the hub, mean, 
and t i p  radius, respectively, was ins ta l led   in  a 14-inch diameter 
single-stage compressor t e s t  r ig ,  The  performance of this blade row 
was ccrmpared over a range of flows at six equivalent t i p  speeds from 
450 t o  915 feet per second for traiUsg-edge thicknesses of 0.045, 
0.030, and O.Ol5 inch. 
AB a check on the results obtained from the campressor teste, a 
brief supplementary Fnvestigation was made on a two-dimemional caacade 
of W A  65-(=)lo blades with trailing-edge thicknesses of 0.045 and 
0.010 inch. 
The following sy~&)ols are wed i n  this report: 
absolute total pressure, (Ib/sq f t )  
radius t o  blade element, ( f t )  
blade-tip  -ve€ocfty,  Ift/sec) 
equivalent tip speed, (ft/sec) 
weight flow, (D/sec) 
weight f lm corrected t o  XACA standEtrd sea-level  conditions, 
( w s e 4  
NACA RM E5lF28 - 3 
6 r a t i o  of +let total   pressure t o  standard  sea-level  pressure 
vd adiabatic  efficiency 
8 r a t i o  of i n l e t  total temperature t o  standard sea-level 
temperature 
Subscripts : 
0 depression task 
2 outlet measuring station 
Comgressor Blading and D e s i g n  
The bLading used for  this  investigation was similar to   tha t  of re- 
ference 6 with the exception of the modified trailing-edge and consisted 
of a single row of 29 rotor blades designed f o r  axial inlet   velocity and 
having a constant solidity of 1.2. Modified W A  65-(14)10, 65-(9.6)09, 
and 65-(7.3108 blower-blade sections were used at hub, m e a n ,  and t i p  sec- 
tions, respectively. Chord length varied from 1.460 a t  the hub to  1.641 
at the mean radius and 1.822 inches a t  the  t ip .  The blades w e r e  fabri6a- 
ted with a trailing-edge thickness of 0.045 inch, which was then reduced 
by hand-finishing t o  0.030 and 0.015 inch, successively. The ra t io  of 
blade trailing-edge thickness to maximum blade thickness is approxhately 
0.10, 0.20, and 0.30 for the 0.015-, 0.030-, and 0.045-inch thicknesses. 
The trailing-edge thickness was  increased equally on the suction 
and pressure surfaces. The theoretical a i r f o i l  was f i r s t  designed with 
zero c d e r  (fig. 1( a) ) and lines were drawn from the  extremities of 
the desired trailing-edge tangent t o  the blade surfaces. The new 
thiclmess  distribution was then  placed on the cambered a b f o i l  
(fig.  l(b) ) . This procedure eliminated the point of inflection that 
exists on the aft portion of the  theoretical   airfoil .  
The blading vas  installed i n  a variable-ccmponent axial-flow com- 
pressor having a constant t i p  diameter of 14 inches and hd-to- t ip  
diameter'ratio of 0.8 a t  the rotor inlet .  Static rotor t ip clearance - was approximately 0.020 inch. 
Compressor Installation  and.Instrmentation 
The compressor FnstaUation i s  s h m  scbmati-cally 3 a  figure 2 and 
is similar t o  that described in reference 6. Instrumentation for 
determination of over-all compreseor performance was sfmil8r t o  that of 
reference 6 and wa8 located a t  the stations indicated i n  figure 2. 
Cascade Installation and B l a d i n g  
The cascade studies were made on a %-inch two-dimensional, solid- 
wall cascade. A set of six NACA 65-(12) 10 blades h a v i n g .  a chmd length 
of 1.5 inches, 8 solidity of 1.9, and a stagger angle of 45' was 
installed i n  the cascade and tested for trailing-edge thicknesses-.of . 
0.010 and 0 .U45 Ini?i"~Thi"%la&a were altered t o  -ob&g  the desired 
trailing-edge thickness in the same manner as the conpressor blades. 
1 
" .  . -- 
For each trai.lin@;-edge thickness, cowressor performance was 
measured a t  equivalent tip speeds of 450, 600, 736, 798, 874, and 
915 feet  per second. A t  each speed, the weight flow varied from 
an approldmate maximum t o  the region of unstable opeation. A constant 
pressure of 25 inches of mercury absolute Uas .haintained.in the depres- 
sion tank for a l l  speeds and weight flm. The range of equivalent 
re la t ive  inlet  Mach n W e r s  and-Reynolds nunibera covered a t  each t i p  
speed is  summrized in   table  I. 
The methods of reference 6 w e r e  used t o  calculate the  total-  
pressure r a t i o  and the adiabatic efficiency of the compressor. Over- 
all total-pressure  ratio was obtained from a mass -fluw-weighted aver- 
age of the  isentrapic energy input integrated 8croas the flaw passage. 
The adiabatic  efficiency was obtained by dividing the integrated 
isentropic energy input to the   a i r  by the  integrated energy input due 
t o  change angular momentum across the  rotor. 
FESWS AM3 DISCUSSION 
A comparison -of blade-raw performance f o r  three bleae trail ing- 
edge thicknesses k t  each of six-equivalent  t ip speeds is sham in  f ig -  
ure 3. The curves-shown w e r e  drawn through the data for the 0.015-inch 
trailing-edge thiclmess. Over the efficient operating range of the 
compressor, that is, t i p  speeds up t o  874 feet  per second, no apprecl- 
able  effect of t r a i u - e d g e .  thickness on total-pres&e r a t i o  was 
measured. The compressor efficiency, however, tended t o  increase wtth 
M C A  RM E51F28 5 
fo 
N 
0, 
w 
increasing trailing-edge thickness. Although the differences (about 1 
t o  2 percent) were wlthin the accuracy of measurement, they w e r e  con- 
s is tent  for  the speeds investigated. A t  an equivalent t i p  speed of 
915 fee t  per second (relative in le t   mch nmiber of about 0.9), an 
appreciable drop in  efficiency occurred, which indicated shock losses. 
Although little change in  total-pressure  ratio  or  efficiency occurred 
between the 0.045- and 0.030-inch thicknesses, a noticeable decrease 
i n  both  total-pressure  ratio and efficiency  took  place for the 
0.0s-inch thickness, which indfcated a lower c r i t i c a l  Mach nrmiber. 
A comparison of these results with those of reference 6 indicates 
that trailing-edge thickness has less effect over the range investigated 
on the performance of a blade row than do o-bher small variations in 
blade row geometry. Although care w&s exercised t o  maintain a constant 
blade angle throughout the investwtion,  a reasonable tolerasce in 
blade angle.was allowed for  the  in i t ia l  se t t ing  of the blades. The 
blade settings for this investigation were probably different from 
those of reference 6 and therefore, although the peak efficiency and 
Pressure ratios were comparable (fig.  3), the weight flows at ,which 
these peak values occurred w e r e  different because of the blade setting. 
In figure 4 a comparison is &de of the radial  variation of total -  
pressure  ratio and adiabatic ef‘ficiency f o r  the three  thiclmesses 
investigated. The curves are plotted for a point i n  the meful oper- 
ating range of the compressor corresponding to a relative  equivalent 
i n l e t  Mach nuiber of approxbately 0.7 and for  a corrected w e i g h t  flow 
of ll pounds per second on the. curves of figure 3( c) . Except next the 
hub and t h e   t i p  where flow conditions are poor, the trailing-edge- 
thickness  effect was negligible even tboughthe chord length and blade 
section vary frm hub t o   t i p .  A t  higher speed6 where a noticeable 
trailing-edge thickness effect occurs the effect, . i n  general,. was the 
same at a l l  radii. 
A rough check on the ccanpressor results with  the two-dimensional 
cascade indicated an increase in l i f t  of about 8 percent when the  blade 
trailing-edge thickness WSLS increased from 0.010 t o  0.045 inch. Wake 
surveys, however, indicated a sufficient broadming of the wake t o  
cause a decrease in lif t-drag ratio of a p p r o ” k l y  4 percent. The 
increase i n  lift is i n  apeement wfth the results obtained in refer- 
ence 5, which ascribe6 the increased lift characteristics for increasing 
trailing-edge thickness t o  the decreasing magnitude and extent of 
adverse pressure gradient over the  t ra i l ing edge of the a f r f o i l .  
The increase in lift caused by the  thicker  trail^ edge could 
result  i n  an increase i n  conpressor efficiency even with a decrease in 
lif t-drag  ratio because the drag measured i n  a cascade includes only 
the  profile losses, which aze a small part of the total losses in a 
compressor. In the congressor tests, however,  no appreciable increase 
i n  l i f t  o r  energy adc2Ltion was observed with  increasing  trailing-edge 
thickness. The basis of this discrepancy  between  the  cascade and the 
6 NACA RM E51F28 
colnpressor results is not evident, b u t  may be the  result of the second- 
ary flows and the effect of rotating-blade wakes on the flow measure- 
ments i n  the compressor. 
The compressor results show that trailing-edge thicknesses up to 
30 percent of maximum blade thickness can be used uithout appreciably 
affecting  the performance of blade sectLons havlng loading d i s t r ibu -  
tions similar t o  that of the NACA 65-series airfoil sections. 
SUMMARY OF RESULTS 
An investigation was made t o  determine some effects of blade 
trailing-edge thickness on the performance of a single-stage axial-flaw 
compressor using a rator row of highly-loaded W A  65-series blower 
blades designed fo r  axial inlet velocity. Over-all compressor perform- 
a c e  was coqtwed far blade trailing-edge thicknesses of 0.945, 0.930, 
and O.Ol5 inch. The trailing-edge thickness w i e d  from 30 t o  10 per- 
cent of maximum b b d e  thickness. A range of weight flm was covered 
a t  each of s ix  equfvalent t i p  speeds Pram 450 t o  915 feet  per second. 
The relative  equivalent  inlet Mach numbers varied from 0.40 t o  0.93 
and the appro"te Reynolds nunibers frm 261,000 t o  604,000. The 
r e s u l t s   a r e   s m r i z e d   a s  follows: 
1. Over the efficient operating speed range of t h e   c q r e s s o r  (LQ 
t o  an equivalent t i p  speed of 874 ft/sec) 110 measurable effect on 
total  pressure was obtained when the blade trailing-edge  thickness wae 
varied f r o m  O.Ol.5 t o  0.945 inch. The efficiency shuwed a tendency t o  
increase slightly with increasing trailing-edge thichess. The mag- 
nitude, however, was within the accuracy of lneasurement (1 to 2 percent). 
2. A t  an equivalent t i p  speed of 935 feet  per second, l i t t l e  
difference i n  total-pressure  ratio and efficiency f o r  the 0.045- and 
0.030-Inch thicknesses occurred. A noticeable decrease i n  performance 
for the 0.015-inch thickness indicated a lower c r i t i c a l  Mach number for 
this  trailing-edge  thickness. . . .. 
3. Trailing-edge thicknesses up t o  30 percent of maximum blade 
thickness were used without sacrifice of performance for NACA 65-series 
blade  s ctions. . .... . . .. . . . . . . . . . . . . . . . .  . " . .. - 
Lewis Flight Propulsion Laboratory, 
Cleveland, Ohio. 
National Advisory Committee for Aeronautice, 
NACA RM E5lF28 7 
l3IPERmm 
1. Reeman, J., and Simonis, E. A. : The Effect of Trailing Edge Thick- 
. 
ness on Blade Loss. Tech. Note No. 116, British R.A.E., March 
1943. 
2. New, Winston R. : Dependence of Stage E f f  fciency upon the  Wailing 
Edge Thichess of the B l a d e s  at  Stibacoustic Velocities. Rep. No. 
N E. lO34, Order No. 6-51002-5 Westinghouse July U, 1939. 
Kt 
f3 3. Sargent, R. F.: T e s t s  on EIACA 0012 wTth Vazious Thick Trailing 
Edges at  High Speeds. R. & M. No. 2209, Brit ish A.R.C., Nov. 
1942. 
4. C a r t e r ,  A. D. S . :  The b w  Speed P e r f o m n c e  of R e l a t e d  Aerofo i l s  i n  
Cascade. Rep. 190. R 55, Brit ish N.G.T.E., Sept. 1949. 
5. Summers, James L., and Page, William A.:  Lift a d  Moment Character- 
i s t i c s  a t  Subsonic Mach Nw&ers of Four 10-Percent-Thick Al r fo i l  
Sections of VaryFng Trailing-Edge Thickness. NACA RM A5OJo9, 
1950. 
6. Dugan, Paul D., Mahoney, John J. , a&. Benser, W i l l i a m  A. : Effect of 
Mach Nu~~iber  on Performance of an Axial-Flow-Compressor Rotor- 
B l a d e  Raw. N X A  RM No. E8D29, 1948. 
a NACA RM E5WZB 
nJ 
w nJ 
0 
NUMBER AT ROTOR IXW3T 
Cquivalent 
; ip speed 
(ft/Gec) 
450 
600 
7 36 
798 
874 
915 
Percent 
Of 
design 
speed 
61.1 
81.5 
100 
108.4 
118.7 
124.3 
T Equivalent relative Reynolds number inlet  Mach nmiber I 
M i n j r m l m  
0.40 
.54 
.66 
.72 
.79 
.81 
I 
aaxirmun Minimum 
0.52 Z.6Wld 
.66 3.59 
.7a 4.37 
.84 
.93 
'5.34 .90 
5.05 
5.71 
3. 56>(105 
4.46 
5.20 
5.52 
5.88 
6.04 
1 
. .  . . . .  . - . . .  
.. . .. , 
. 
. . . .  Id 
B 
I I I I I I I I I I I 
0 .I .2  .3 .4 .s .6 .7 .a .9 1.0 
Distance along chord, fraction of chord 
(b) Final a i r f o i l  shape. 
Fimre  1. - A l r p o l L  w o f i l e  at mean radius for  three traiLlm-e&te thicknesses. 
. .  
. .  
.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -. . . " .  . . .. .. - .  . .  
moe tmk 7 Inlet t b r o t t l e  7 
Outlet, 
S t a t I a n  e --, 
Gutlet 
tl!nlttla - 
. .  . .  
4 
. . .  ... . . . . . . . . . . 
I I 
. . . .. 
. 
Corrsotad aaight 
I Corrected weight 
( 0 )  Equivslaat tl~/sperd, 756 feat per  m o o d .  
1 * 
. .  . " 
. . . . . . . . " ..  .. .. . 
4 
. . . . . . . . . . . . . . . . .  . . . .. 
b 
.. . .. 
* 
I 
(.) Eqniwlmt t i p  1p8d, 874 fsbt per a m i d .  
14 NACA RM E5lF28 
". 
,. 3 
..2 
.80 .&4 .96 1.00 . **. . . . . . -. .~ 92 
Radius ratio,  r/rt T 
Figure 4. - Radial variation of total-pressure  ratio and 
adiabatic efficiency for three trailing-edge thickneeses at 
corrected weight flow of 11 pounds per second for equivalent 
tip speed of. 736 feet per second. 
NACA-LanelW - 10-28-51 - SOQ 
- . . . . . . . . . .  - - -  . . . . . . . . . . .  
- - .  ....... 
-.. - .  - - -  
. . . . . . . . . . . . . . .  :- 
.~.... " . i. - ..  ". - "" 
" ............ .,.. . .: ." .C . 


Some effects of blade trailing-edge thickness on performance of a single-stage axial-flow compressor

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server