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NASA Contractor Report 198362
4- 93
Off-Design Computer Code for Calculating the Aerodynamic Performance of Axial-Flow Fans and Compressors User's Manual
James E Schmidt NYMA, Inc. Brook Park, Ohio
(NASA-CR-198362) OFF-DESIGN N96-12502
COluiQUTER CODE FOR CALCULATING THE AERODYNAHIC PERFORMANCE OF AXIAL-FLOW FANS AND COWRESSORS Unclas Final Report <NYMA) 9 3 p
6 3 / 0 7 00605,a7
July 1995
Prepared for Lewis Research Center Under Contract NAS3-27186
National Aeronautics and Space Administration
https://ntrs.nasa.gov/search.jsp?R=19960002494 2018-01-31T19:58:59+00:00Z
TABLE OF CONTENTS
Page .
summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1
2.0 Off-Design Compressor Calculation Procedure
3.0 Input Data . . . . . . . . . . . . . . . . . . . . . . . . . . 5
3.1 General Input . . . . . . . . . . . . . . . . . . . . . . 5
4.0 Aerodynamic Flow Solution . . . . . . . . . . . . . . . . . . 9
1.0 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . 2
. . . . . . . . . . 4
3.2 Calculation Station and Blade Row Input . . . . . . . . . . 6
4.1 Flow Loss and Deviation Prediction Models . . . . . . . . . 9
5.0 Exit Calculation Routine . . . . . . . . . . . . . . . . . . . 11
6.0 Controlling Recalculation Main Routine . . . . . . . . . . . . 12
7.0 User Example Compressor Case . . . . . . . . . . . . . . . . 13
8.0 Acknowledgement . . . . . . . . . . . . . . . . . . . . . . 15
9.0 References . . . . . . . . . . . . . . . . . . . . . . . . . 16
Appendix
ASymbols . . . . . . . . . . . . . . . . . . . . . . . . . . 17
B Input Parameters for Off-Design Compressor Code . . . . . . . . 18
Tables
1 Example Input Data Set (no Stator Resets) . . . . . . . . . . 32
2 Example Output Data Set (no Stator Resets) . . . . . . . . . 36
3 Example Output Data Set (with Stator Resets) . . . . . . . . 64
iii
SUMMARY
An off-design axial-flow compressor code is presented and is available from
COSMIC for predicting the aerodynamic performance maps of fans and com-
pressors. Steady axisymmetric flow is assumed and the aerodynamic solution
reduces to solving the two-dimensional flow field in the meridional plane. A
streamline curvature method is used for calculating this flow-field outside the
blade rows. This code allows for bleed flows and the first five stators can be reset
for each rotational speed; capabiIities which are necessary for large multi-stage
compressors. The accuracy of the off-design performance predictions depend
upon the validity of the flow loss and deviation correlation models. These em-
pirical correlations for the flow loss and deviation are used to model the real flow
effects and the off-design code will compute through small reverse flow regions.
The input to this off-design code is fully described and a user’s example
case for a two-stage fan is included with complete input and output data sets.
Also, a-comparison of the off-design code predictions with experimental data is
included which genedy shows good agreement.
1
1.0 INTRODUCTION
If a reliable off-design axial compressor code is available for predicting the
performance maps of compressor designs, costly and time consuming testing can
be kept to a minimum. In addition, reliable flow property prediction at off-design
speeds can be used as a good starting point for full three dimensional flow anal-
yses to reduce their long computing times. off-design compressor performance,
which is the performance of a compressor at flow conditions and rotational speeds
other than the design point is an important aspect of compressor design. A good
design for a multistage axial compressor requires an adequate part-speed per-
formance with s&cient part-speed stall margin and determination of required
starting bleed flows along with any guide w e or stator resets needed to better
match the stages (Ref. 1). The off-design flow analysis is different from the
compressor design analysis in that the compressor blade geometry is now &ed
and the off-design requirement is to determine the compressor performance for
a range--of speeds and weight flows i.e. to determine the compressor map.
This off-design, two-dimensional, axial compressor code was originally de-
veloped by James Crouse as an analysis code to be used in tandem with his
compressor design code (Ref. 2). This off-design code has never been published
and many modifications to the original code have been made over the last 12
years. In order to successfully use this off-design code, the compressor design
code must first be used to generate the necessary blade geometry block of data
that is part of the required input to the off-design code. Actually, the general
design procedure is to first produce a preliminary initial blade design with the
compressor design code. In cases where the compressor blade design is obtained
from an old design or an outside source, the compressor design code must be used
to duplicate the blade design and obtain the necessary input to the off-design
code.
Steady, axisymmetric, two-dimensional flow in the meridional plane is as-
sumed and the equations of motion as developed in reference 1 are only applica-
ble for calculation stations outside the blade rows, Also, empirical correlations
for the flow losses and deviation angle are used to model real flow effects and
the off-design code will compute through small reverse flow regions.
The blade geometry desaiption is essentially the same as that for the com-
pressor design code (Ref. 2) and the aerodynamic solution uses a streamline
curvature method for calculating the flow at stations outside the blade rows.
The aerodynamic calculation structure for the off-design code is set up such
that the aerodynamic flow field and selected streamlines are recalculated for
each new rotational speed with the flow rate increased to the choke condition,
and then decreased to the s t d condition. In addition, this code has the capa-
bility to include bleed flows as well as reset adjustments for the first five stators.
The purpose of this report was not intended to give a detailed description of
the development of the off-design code but to give the user a manual for properly
using this code.
A user’s example case for a two-stage fan is included to demonstrate the
necessary input required to run this off-design code as well as the resulting
output produced by this code. A comparison of the off-design code performance
predictions with experimental data for this two- stage fan is made with and
without stator resets.
3
2.0 OFF-DESIGN CALCULATION PROCEDURE
Similar to the compressor design code (Ref. 2) the off-design code is sep-
arated into three main regions of calculation: the input and setup, the aero-
dynamic flow solution and exit calculation routines. In addition, a reset or
recalculation routine is included for a recalculation of the flow field and stream-
lines as the flow is increased toward the choke point and decreased toward the
stall condition. This general procedure is repeated for each desired rotational
speed.
For the input and setup routine the input data are read and interpreted
and the streamlines are located on the basis of annulus area. For the flow field
solution phase the equations of motion are satisiied in the meridional (r-z) plane
at calculation stations along the compressor flow path (see Figure 1). At the
stations, the equations of motion and overall flow continuity are satisfied using a
streamline curvature method of solution. The exit calculation routine makes the
f2m.I ciilculations and prints the output. First the mass averaged parameters for
the individual and cumulative compressor blade rows are calculated and printed
and then the detailed tabulated local dues (tip to hub) of the aerodynamic
and blading parameters for each calculation stations are computed and printed.
A recalculation phase or controlling main routine is used to recalculate the flow
field and streamlines as the inlet flow rate is increased toward the choke point
and decreased toward the s t d condition. This procedure is repeated for each
desired compressor rotational speed.
r \ x r ~ o f ~
F@UR L -'GalcuiatiW, stations in m m r flow pah
4
3.0 INPUT DATA
The basic meridional ( r , z ) computational plane is shown in Figure 1 for
a typical compressor configuration. As indicated in Figure 1 the calculation
stations are located at the blade-row leading and trailing edges and at other
annular locations needed for proper determination of the streamlines. The input
data are divided into two categories of input: general input data and calculation
station and blade row input data. All the input parameters and options as well as
the input data format are described in Appendix B. The mathematical symbols
used for defining these input parameters are given in Appendix A. Except for a
few modifications the general input data for the off-design code are similar to
the general input data for the design code (Ref. 2).
3.1 General Input Data
The input data consist of the following information:
1). N-ber of desired compressor speeds.
2). Normalized range for the choke region.
3). Normalized step increment for the choke region.
4). Normalized range for the stall region.
5). Normalized step increment for the stall region.
6). Fkaction of desired rotation speeds.
7). Stator reset angle for first 5 stators.
8). Number of streamlines.
9). Number of blade rows.
10). Number of annular stations.
11). Compressor rotational speed.
12). Inlet flow rate.
13). Desired compressor pressure ratio.
14). Gas molecular weight.
15). Co&cients for Cp as a Hth-degree polynomial function of temperature.
5
16). Upstream values of total temperature, total pressure, and inlet tangential
velocity for each streamline.
17). Streamtube m a s flow fractions between streamlines.
18). Sets of points to define tip and hub casing contours.
19). Sets of blade element profile loss parameters that are tabulated as func-
tions of blade element loading parameter (D Factor) and fraction of passage
height.
Except for the above first seven input data items the remaining list of the
general input data is identical to the general input data for the design code
(Ref. 2). Therefore, once the design code is run most of the general input data
for the off-design code axe already known.
3.2 Calculation Station and Blade Row Input
The input data for the annulus CalcuIation stations and blade rows are read
in ordq from annulus inlet to outlet. The calculation stations and blade row
input consist of the following information:
I. Type of calculation station: annular, rotor or stator.
If annular - only one line.
1). The tip and hub axial locations.
2). The tip and hub blockage displacement thickness.
3). The cumulative bleed fraction of inlet flow
If rotor - h t line.
1). The tip and hub stacking-point axial locations.
2). The inlet tip and hub blockage displacement thicknesses.
3). The cumulative bleed fraction of inlet flow.
If rotor - second line.
1). Number of streamlines (maximum of 11).
2). Loss set number (maximum of 9).
6
3). Blade d&tion index.
4). Option parameter for part span shroud (IDAMP = 0 for no damper).
5) . The cumulative bleed factor of inlet flow.
6). Number of blades.
7). The outlet tip and hub blockage displacement thicknesses.
8). The angle of stacking axis tilt (degrees) or lean in circumferential direction
( r , 6 plane).
9). The estimated cumulative stage adiabatic efficiency for the first iteration.
10). Desired cumulative energy addition factor through particular rotor to total
energy addition of compressor. The last rotor must have CRENGY = 1.0
to meet the input total pressure ratio.
If stator - first line.
1). The tip and hub stacking-point axial location.
2). The inlet tip and hub blockage displacement thicknesses.
3). The cumulative bleed fraction of inlet flow.
If stator - second line.
1). Number of streamlines (maximum of 11).
2). Loss set number (maximum of 9).
3). Blade deibition index.
4). Option parameter for part span shroud (IDAMP = 0 for no damper).
5). The cumulative bleed &action of inlet flow.
6). Number of blades.
7). The outlet tip and hub blockage displacement thicknesses.
If IDAMP > 0, then the following information line needs to be included
after the second line of the rotor or stator input.
1). Radial location of the leading edge or the part-span shroud.
2). Slope angle of the damper centerline in the meridional view.
7
3). Chord length of mid-span shroud.
4). Maximum thickness-to-chord ratio of damper.
5). Leading edge to chord ratio of a part-span shroud.
6). Tkailing edge to chord ratio of a part-span shroud.
The basic blade row geometry block of information is included after the
second line of the rotor or stator blade row input. This basic blade row geometry
data is automatically obtained using an option in the compressor design code
(Ref. 1). Again, as previously stated, the order and format for all of the off-
design input is shown in Appendix B along with a complete description of the
input parameters.
8
4.0 AERODYNAMIC FLOW SOLUTION
The aerodynamic flow solution determines the velocity diagrams and fluid
state properties on the calculated streamlines at the annular and blade row inlet
and exit stations. The flow solution results from an inner and outer iteration
scheme. For the outer loop iteration the calculated variables are the stagna-
tion temperature and pressure, the tangential component of velocity, and the
streamline location, slope, and curvature, The inner loop iteration is the span-
wise flow calcdation, solving continuity and radial equilibrium at each station,
where the axial velocity distribution is calculated as the outer loop variables are
held fixed and then the streamline location is reset. The o v e d procedure is
station marching from inlet to outlet with streamline parameters fixed and a
maximum of 20 cycles to converge to a solution.
4.1 Flow Loss and Deviation Prediction Models
Accuracy of the off-design performance predictions depends upon the valid-
ity of the flow loss and deviation correlation models. The total loss coefficient
is defined as:
- (P0f)Z -Po; W =
Po: -P1
In this code the losses consist of two additive components: shock losses and
all other losses (considered as a proflle loss). The shock loss is assumed to be
approximately equal to the computed normal shock loss divided by the square
of the inlet relative Mach number. The proflle loss coefficient is given by:
- wp =F-F*
where ga is the shock loss Coefficient.
The profile loss is represented by a correlation parameter as a function of
passage height and aerodynamic blade loading. The values for this loss parame
ter correlation axe input in tabular form in the input data and axe usually called
9
loss parameter data sets. The aerodynamic blade loading parameter (more com-
monly called D-factor) is de&ed as:
where Ve is the tangential flow velocity and u is the blade element solidity and
the loss parameter in the input table is deiined as:
Gp cos p; 2u (4)
Loss parameter data sets determined from experimental data over many
years are often stored in data banks for different types of compressors.
In this off-design code a deviation subroutine is used to iterate for the
blade row outlet flow angle and thus with fixed blade geometry determines the
deviation angle. A correction to Carter's deviation angle nile (Ref. 1) due to
changes in blade camber in the s t readhe direction was developed by James
Cmuse in his original developrnent of the off-design code. This correlation for
the deviation coetficient used in the iteration for the outlet blade flow angle is
given by:
C, = (0.219 + 7(0.051085 + 0.0889157))
x (1.620 + ~(0.72365 - 0.278717)
+ qi[-1.415 + ~(1.85240 - 0.75964-/)
+ qi(0.352 - 7(0.810816 - 0.4044ey))]}2"
where: 7 is the blade setting angle (rad.), c is the chord, and 4 is the ratio of
front to total cambc; of each blade element.
These loss and deviation coacient correlations are the primary unknown factors in the off- design code. Generally, the off-design code predictions com-
pare quite well with experimental data for fans and lightly to moderately loaded
10
multi-stage compressors. However, for highly-loaded, large multistage compres-
sors very large reset guide vane and stator adjustments are needed to match the
stages, especially at low wheel speeds. These large reset adjustments greatly
increase the pressure losses and deviation angles in these front stators and guide
vane. Therefore, the loss and deviation correlations need to be improved for
highly-loaded, large, multistage compressors.
5.0 EXIT CALCULATION ROUTINE
After the iterated solution for the aerodynamic flow field is obtained the
off-design code calculates and prints the overall mass-averaged parameters for
the individual and cumulative compressor blade rows. In addition, the detailed
tabulated local values (tip to hub) of the aerodynamic and blading parameters
for each calculation station are computed and printed.
11
6.0 CONTROLLING RECALCULATION MAIN ROUTINE
The first flow field and performance calculation for this off-design code is
the design point case and the results should compare very well with the results
for the design code (Ref. 2). Then, the step-choke increment is used to calculate
an increased flow rate and the flow field with new streamlines are recalculated.
This process is repeated until the input range-choke value is reached and all the
performance calculations on the choke side are completed. A word of caution
is needed at this point because if too large a value for the input range-choke is
used, the code will try to calculate a very hard choke condition which may cause
a divide by almost zero and other numerical difficulties. After the choke-side
calculation is completed, the off-design code then uses the step-stall increment
to reduce the flow rate from the design value and the flow field with new stream-
lines are calculated. This process on the stall side is repeated until the desired
input range-stall value is reached. Again, caution is needed because if too large
a value'-for the input range-stall is used, the code will try to calculate a stall
condition with very high D-factors (Eq. 3) resulting in severe numerical diffi-
culties which can stop the code. Now, all the flow-field overall and blade row
performance calculations have been obtained for one complete speedline (100
percent or fraction of speed of 1.0). This same general procedure is repeated for
additional speedlines except that instead of marching away from design point
values the controlling main routine calculates a reduced flow rate and rotational
speed for each additional desired input fraction of speed. Speedlines greater
than 100 percent or fraction of speed greater than 1.0 can also be calculated.
12
7.0 USER EXAMPLE COMPRESSOR CASE
A user’s example case for a two-stage fan (Ref. 3) shows the necessary
detailed input data set needed to run this off-design code (see Table 1). This
twestage fan example case is the same example case that was used for the
compressor design code (Ref. 2). In addition, the off-design output prediction
results for this example case with and without stator resets are shown in Table 2
and Table 3, respectively. A comparison of the off-design predictions is generally
in good agreement with the experimental data (see Fig. 2), especially for the 70
percent speed case.
This compressor off-design computer code (written in Fortran) dong with
the example input case can be obtained from COSMIC, 112 Barrow Hall, Uni- versity of Georgia, 30601. The COSMIC program number is LEW-16176.
13
LOW-ASPECT RATIO TWO STAGE FAN(67) off-design cal.(no reset) off-design cal.(with reset) - - - .
2.8
2.6
2.4
2.2 0
.w CI,
2 2.0 d) k a * 1.8
pc
1.6
1.4
1.2
1 .o
I I
I I I I
; I I
..... f - ........... I: > ....... I I I I
30 40 50 60 70 80
Flow, (lbdsec)
Figure 2 - Comparison of the off-design code perfomance predictions with experimental data for a low-aspect-ratio two-stage fan with and without stator reset.
14
8.0 ACKNOWLEDGEMENT
The original version of this off-design code was developed by James Crouse
while he was an employee of NASA. Even after he left NASA James Crouse
has provided several corrections and additions as well a5 many suggestions to
improve this code and his valuable help is greatly appreciated.
15
9.0 REFERENCES
1. Johnson, I.A., and Bullock, R.O. eds., “Aerodynamic Design of Axial-Flow Compressors”, NASA SP-36, 1965.
2. Crow, J.E., “Computer Program for Aerodynamic and Blading Design of Multi- Stage Axial-Flow Compressors”, NASA TP-1946,1981.
3. Urasek, D.C., Gorrell, W.T., and Cunnam, W.S., “Performance of Two- Stage Fan Having Low-Aspect Ratio, First-Stage Rotor Blading”, NASA TP-1493, 1979.
16
APPENDIX A
Symbols
A C
C,
D I i J K k L M m
P PO
R
S TO t
U V W
f
z
annulus area, ft2 blade chord, in. specific heat function for constant pressure, ft/seca OR empirical correlation for the deviation angle coefficient blade element diffusion factor integer station index incidence angle, deg. integer streamline index ineger loss set index curvature in curvilinear cc-ordinate system distance along chord l i e , in. Mach number streamline direction in meridional plane, in. static pressure, lb/ft2 stagnation pressure, lb/ft2; input values; lb/i2 gas constant for air, ft-lb/slug O R
radius from axis of rotation, in. blade element path distance, in. stagnation temperature, O R
static temperature, O R ; also, blade element thickness, in. local blade velocity, ft/sec. velocity, filsec. weight flow, ft/sec. axial distance, in.
0
P
Y 6 B 4 IC
P Q
i3 ws - rj,
Subscripts: I i le m
0
t t e B 1 2
Superscripts: ( Y
angle of streamline with reference to axial direction, deg. flow angle relative to meridional direction, deg. blade setting angle, deg. deviation angle, deg. circumferential direction, rad ratio of front to total camber blade angle relative *, local conic co-ordinate ray, deg. local density, slugs/ft3 local blade solidity, ratio of chord to circumferential spacing total loss co&cient shock loss coefficient profile loss coefficient
calculation station index ideal value, as by an isentropic process leading edge streamline dirction in meridional plane; also, maximurn blade thickness initial value transition point trailiig edge circumferential direction blade row inlet blade row outlet
relative to rotor
17
18
I
19
20
x E:
.d 8
&
.e s
-9 d
3 M 9
I 8
- u la
u
3
-8 s
W
d
v m X E V
+ n m m X E V V +
h
v 3
E V 4
2 E
d m X d V
i- m m X d V V + m X d u Q + m X
n
N
2 F
x E
v m X
i2 n 3
21
d m X d u +
m m X
n
g u + N m X d V D
.- + m X
dl m X E E+
I m X
V I
m X E E-c D
I
n m
2 N
*?
@E +
d m X
E n I
m X d E V
I m X
m
N
d
I n
!s @Is
dl m X Fr V R + 0 m X F V V + m X Fr V D + m X
N
w m X
E.3
E3
E.3
n + m X
m
V + m X
N
E9 + m X
22
-=?
F 2
* m X d
- h
v 3 F
V n 8 V
.r m X F V
+ m X E V V + m X E V
+
n 0
N
m m X
23
2 E
.r m X E E
I m X F
I cn X E E
I
n m
F1
m
.r m X
F2 n I
m m X p: E V
I m X
F1
F2 m
x E:
x E:
24
r 4 - I
I . : P
< P P <
E i; 2
€
t i e
2 C
P
i c
+
6
.e 8 4 0 .- 3 b a
I
2
+a
B d
PI
c3 Y
n 0 0 4
W 0
8 'S E
E v
"cr d d 0 d ca d
d
.-
... 4
4
.41
-Ea V
+a
3
Y
v3 Y
2s
I I I
26
5 s E!
c E 7
.- e 6 s
d w E
hi a h c
.e f3 4
e d
27
‘tt
k 5:
a 41 fi 9
6
7
a Y
a d 0 0
w
.- Y
d 8
-5
gb z 3 d
a
0 Y
bo E
GI
5 62
‘r
E 2
a 41 4 9 Pe P
0
a Y
a d 0 0
W
.e Y d 8
-5
8 z
P
a
0 Y
bo E L=: -2
w E 62
28
h Y z. a
m -- g
-9 8
CS P 8
&
Y
4 LI
Y 0 d
3; n
M W a
3 T s m
29
“i E: E
s E B 3 e,
a 63
w
0 .- Y E z 4 0 m
a
Y
-2 d 8b z 3
3
2 n 3 W
X H
‘11
E E:
“i
E E:
30
E
I c C .. + .. l
E
c *
i! d
n a
31
x :: x x x x x 9xx x x x x 2x88 x x 8 x x x x x x x x x 8 x x x x x x x x x o e C. o o o o o m m 0 0 0 0 o o o c o o o o o o o o o o o o o o o o o o o o o o
0 0 0 0 m N I I 0 0 0 0 0 0 0 4 " " " " F F F " ~ " " C . ~ 7 ~ F " 0 . m m *
. e - c w o . . . W ? ? m " ? ? . . ym'Dq * N O O m m o 1 1 m - 0 0 1 1 - w 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 In 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0
r l N rl I d N
- B U P O D - e 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 w w w w w w w w w w w w w w w w ~ ~ ~ w ~ w g 3x2 zog 0 0 0 0 m o o 0 . . . . . . . . . . . . . . . . . . . .
(0 v) m m m m o o - *;;id o o o o o o o o o o o o o o o o o o o o o o
y L D m v ) v ) L D m m Y ) m v ) y m m v ) m m L D L D m m m
G: x ? : 0 0 0 - N O N 0 m C. 2 II
N 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0
O W L D O 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 - 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 1 1 0 e
- t m w ? ? y ecr-? . . m m ~ e I -4
. . . . . . . . . . . . . . . . . . . . O O O O O C o o o o o c o o o o o c
m r l o m - 4 my"? W W W N O O O ~ ~ O O O O O O O O O ~ O O O O ~ ~ ~ ~ ~ ~ ~ ~ ~
m m m m m ~ ~ m m n w y 1 1 o O O O O O O O O O O O O O ~ O O O ~ O O ~ ~ ~ ~ ~ ~ ~ ~
o m mo I I O O ?
~ m m Y) 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 . . . . . . . . . . . . m m r( rl
4 II II 4 4 4 4 II. e 4 4 4 4 4 4 4 4 4 II. 4 4 4 . . I -
0 0 0 0 0 0 0 0 0 0
0 0 - 0 0 0 0 0 0 0
r ( c w o m - rl m w I d s 4 ..I ..I
0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0
::xs 8x2 P l t Z -4:- ? ? ? ? . .
a o o o o o o m 0 0 ? N ? -!?? N N ? ? m e i n o m - N - ~ I I
8 4 - 4 , -
0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 orlo 0 0 0 D O D O
c m o o m m O N W ~
O W ? . . q ? ? : I 4 . I - 8 -
0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 O o O O O O O n O O O O O 0 0 0 0 0 0 0 0 0 1 0 0 0
I I I I - N O ~ ~ - I I ~ W N N ????-!?????-!?? r( - N r l -..In - N
32
0
0 Y
0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0
0 0 D O 0 0 0 0 0 0 0 0 0 0
0 0 0 0 0 0 0 0 0 0 0 0 8:: 2:: 8x x x EX 85: P O P P 0 0 0 0 0 0 0 0 0 0 0 0 0 0 x x gg 80" x x 0 0 . . 9 9 9 9 9 9 9 9 33 9 9 9 9 9 9 9 9 39
33
k B H
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REPORT DOCUMENTATION PAGE
1. AGENCY USE ONLY (Leaveblank)
Form Approved OM6 No. 0704-0188
2. REPORT DATE 3. REPORTTYPE AND DATESCOVERED
July 1995 Final Contractor Report 4. TITLE AND SUBTITLE
Off-Design Computer Code for Calculating the Aerodynamic Performance of Axial-Flow Fans and Compressors User's Manual
6. AUTHORtS)
James E Schmidt I AND ADDRESS(ES)
NYMA, Inc. 2001 Aerospace Parkway Brook Park, Ohio 44142
5. FUNDING NUMBERS
WU-538-03-11 C-NAS3-27186
8. PERFORMING ORGANIZATION REPORT NUMBER
E9764
12a. DISTRIBUTtOWAVAILABLITY STATEMEM
I 9. SPONSORIN~ONITORNG AGENCY NAME@) AND ADDRESS(ES) 10. SPONSORINWMONITORING
AGENCY REPORT NUMBER
12b. DISCRIBLITION CODE
National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 44135-3191
14. SUBJECTTERMS
Off-design aerodynamic performance; Axial-Flow fans; Multi-stage axial compressors
NASA CR-198362
15. NUMBER OF PAGES
94 15. pRcE CODE
17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFlCAl'lON OF REPORT OFTHIS PAGE OFABSTRACT
A05 20. LIMITATION OF ABSTRACT
I I I Unclassified Unclassified UncIasdied I NSN 7540-01-2805500 Standard Form 298 (Rev. 2-89)
Presorbed by ANSI Std. 23918 298-102
