Efforts continue to investigate the applicability of the Deiwert's time dependent numerical airfoil calculation code to the simulation of two dimensional airfoil flows with large amounts of separation present. A modified logarithmic region mixing length term produced substantial differences in the calculated flow fields for the large separation of the M=0.4, high angle of attack cases. Reasons are given for interest in applying that model to previously investigated cases of the 64A010 airfoil section at Mach number 0.8 for angles of attack of 4 deg and 6.2 deg. The effects of these turbulence models at alterations and the time dependent behavior from impulsively started conditions are discussed

Rose, W. C.

English

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https://ntrs.nasa.gov/search.jsp?R=19830010391 2020-03-21T05:39:53+00:00Z
(YA4A-CR-•169927)	 COMPUTATIONS AND TURHULFNT	 N83- 18',62
FLOW MODELING IN SUPPORT nF HK ICOFTER ROTOR
TECF,NO ►_OGY Progress Report, 1 Jun. - 30
Nov. 1982 ;Nevada Univ.)	 17 p HC A02/MF 401	 UllC1aS
CSCL 01A G3/02 02863
ENGINEERING RESEARCH
AND DEVELOPMENT CENTER
COLLEGE OF ENGINEERING
UNIVERSITY OF NEVADA, RENO
RENO, NEVADA 89557
(702) 784-6925
	
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TECES-LOGT Progress report. 1 Jan. - 3^
l3v. S S82 ;levada Univ.) 17 D !K 102/9? 411	 Onelas
CSCL 11t 63/02 Zj2663
ENGINEERING RESEARCH
AND DEVELOPMENT CENTER
COLLEGE OF ENGINEERING
UNIVERSITY OF NEVADA, RENO
RENO, NEVADA 89557
(702) 784-6925
ii
J
P
'ROGRESS REPORT
JUNE 1, 1982 THROUGH NOVEMBER 30, 1982
NASA GRANT NO_ NSG 2291
1
e
COMPUTATIONS AND TURBULENT FLOW MODELING IN SUPPORT
OF HELICOPTER ROTOR TECHNOLOGY
Prepared for:
Aerodynamic Research Branch
NASA-Ames Research Center
Moffett Field, CA 94035
by
i
William C. Rose
Engineering Research and Development Center
University of Nevada, Reno
Reno, NV 89557
PROGRESS REPORT
JMv_ 1, 19x.:2 THROUGH NOVEMBER 30, 1982
BACKGROUND
Continuing efforts on the grant were expended to investigate
the feasibility of using Deiwert's time-dependent numerical
simulation code to calculate two-dimensional airfoil flows
similar to those expected to occur on helicopter rotors. Rather
than attempting to calculate the actual helicopter oscillating,
unsteady flow, critical components thought to be important to
the overall f?ow are being studied individually. Since the
helicopter rotor problem involves airfoil sections at large
local angles of attack, both time accuracy and adequate turbu-
lence models are required.
IEfforts during the present reporting period continued
to investigate the applicability of the Deiwert's time-dependent
numerical airfoil calculation code to the simulation of two-
dimensional airfoil flows with large amounts of separatica pre-
sent. In the previous Progress Report for the period December 1,
1981 through May 31, 1482, it was shown that at angles of attack
of 12.5 to 17.5 deg at a Mach number of 0.4, non-physical behavior
of a flow was observed. This was a manifest in a nearly periodic
shedding of the entire flow field. In that report, a variation
in the wall region turbulence modeling was proposed. The usual
wall layer is modeled with the Karr^an constant as:
k = ky
	 (1)
-1-
t
t
I	 A model. which has a decreasing-wall influence in separationsg	 P
I
was proposed as equation 2.
I = ky(10 log (100 (ysep ys)/c)j	 (2)
The net effect of Equation 2 is to reduce the Karman con-
stant whenever the height of the separation bubble relative to
the chord length becomes significant. If, for example, the sepa-
ration height is approximately 10% of the chord, the Karman con-
stant is increased by a factor of 10. The calculations carried
out in the previous progress period had indicated a rather sub-
stantial change in cne calculated flow properties, particularly
over the aft portion of the airfoil. Extensive separation was
calculated and velocity profiles similar to those observed
experimentally were predicted.
Modification of the Karman constant as proposed during
last period's reporting has been one of the uncountable number
of modifications to turbulence models which might be proposed.
For purposes of discussion here, the boundary layer on an air-
foil undergoing a severe adverse pressure gradient leading to
separation can be divided into three principle regions. First
the outmost layer, or so-called wake region, in which the eddy
viscosity or mixing length is held approximately constant. The
logarithmic region of the velocity profile, which is described
usually by Equation 1, which decreases the mixing length or
eddy viscosity to near zero at the wall, and finally an inner-
wall region which is usually described by a Van Driest damping
-2-
term in order to account for the existence of a molecular sub-
layer. Clearly the modification discussed in last period's
progress report is a modification to the logarithmic region of
the boundary layer. It was shown there that minor modifications
to this region of the flow produce extensive changes in the
calculated flow-field behavior. With respect to the outer
region, extensive _!xperiments in the NASA 2'x2' tunnel in making
direct measurements of a Reynolds shear stress term with laser
velocimetry has indicated that a very good approximation to
the outer layer behavior is given by the mixing length term,
.096. This is true even for flows having just undergone abrupt
adverse pressure gradients resulting from shock waves (including
separated flow). Thus, the suspected non-equalibrium nature
of the outer portion of the flow does not seem to be verified
from experiment and that the disagreement between the calculated
flow fields and experimental flow fields cannot be blamed on
outer layer Reynolds shear stress behavior alone. The modified
logarithmic region mixing length term, as noted above, did
produce substantial differences in the calculated flow fields
for the large separation on the M=0.4, high angle of attack
cases.
In this report, that same model has been applied to
the previously investigated cases of the 64AO10 airfoil section
at Mach number 0.8 for angles of attack of 4 deg and 6.2 deg.
The interest in these two cases lies in the following:
-3-
is
'.r
r
i
Although both flow fields contain some separated flow, the
flow at the 6.2 deg angle of attack does have a substantial
region of backflow, experimentally, and has not been properly
calculated in the past using the Escudier model throughout
the course of the present study. The 4 deg case contains
only a minor amount of separation and it is not expected that
the alteration of a logarithmic profile in this particular
case should have a substantial influence on the calculated
solution. The interest in the 4 deg case at this point lies
in the unsteady nature of the solutions obtained using Steger's
code by King, Degani and Chyu and not obtained previously with
the Deiwert code for the 4 deg case. Effects of these turbu-
lence model alterations and the time dependent behavior from
impulsively started conditions are discussed in the present
report.
-4-
l
3
f
RESULTS AND DISCUSSION
in order to illustrate the behavior of the 64A010 airfoil
section at 4 deg angle of attack and Mach number 0.8, the
first portion of this study was carried out. Figure 1 shows
the upper and lower surface airfoil pressure distributions
obtained from the Deiwert code with the Escudier turbulence
model for the 4 deg angle of attack condition. Free-air
boundary conditions are used throughout this report even though
the free air assumption is known to be grossly inadequate in
the NASA Ames 2 1 x2' wind tunnel at M=0.8. These boundary
conditions are uniformly duplicable in all of the analytical tech-
niques and are thus chosen here for illustration. The airfoil
pressure distributions shown in Figure 1 were obtained after
approximately 25 chords of calculation from an impulsive start
to the Mach 0.8 condition. Reynolds numbers in this report are
2.0x106
 based on chord length unless otherwise noted specifically
in the discussion or on the figures. The mean average lift
coefficient for the 4 deg case calculated from the present theory
has a value of approximately 0.72. The time evolution of this
 lift coefficient is shown in Figure 2 where a monotonic progression
from 0 to a final, constant value of approximately 0.72 is seen.
Over approximately the last ' 10 chords of calculation time, the
lift coefficient is steady and indicates no trend whatsoever
toward the periodic behavior as obtained by others using the
Steger code. The effect of changing from the Escudier model
-5-
to the wall model of Equation 2 is shown in the upper and lower
surface pressure distributions at 4 deg and Mach 0.8 in Figure 3.
Because only a small amount of separation is present in the
calculation for this case, no large differences are indicated
between these two turbulence models. Again, the progression
from the impulsive start to the condition shown at approximately
25 chords of flow exhibited the exact behavior as that shown
for the Escudier model in Figure 2. Unsteady periodic behavior
has also been observed in other calculations published recently
in the April 1982 AIAA Journal by Sugavanam and Wu. The fact
that the Deiwert code does not exhibit this periodic behavior
for this type of solution at the lower angles of attack may be 	
I
due to the use of the MacCormack algorithm as opposed to the
Steger algorithm for the solution of the equations. Additional
numerical damping inherent in the MacCormack algorithm may be
one possible explanation for the lack of calculated periodicity
in the flow fields, although this point remains speculative.
In order to investigate the effect of changing turbulence
models for a case in which extensive separation is expected to
occur, the following series of runs was performed for the 64AO10
section at a Mach number of 0.8, again for the free air boundary
conditions, at an angle of attack of 6.2 deg. Figure 4 shows
a comparison between the final steady upper and lower surface
pressure distributions obtained for this flow using the Escudier
and modified wall models. The steady lift coefficient for the
-6-
Escudier model is approximately 0.78 while that calculated from
the modified wall model is approximately 0.94. Unfortunately, the
modified wall model predicts a flow that remains attached further
than that of the Escudier model. This is opposite to the behavior
1
expected based on the calculations performed during the last
reporting period on the 0012 section at M=:.4, far which a
much more extensive region of separated flow was calculated
with the modified wall turbulence model.
The time history of the lift coefficients for successive
changes from Escudier to the wall model are shown in Figure S.
Approximately 25 chords of flow from the impulsive start were
calculated with the modified wall model and the turbulence model
was changed instantaneously to the Escudier model. A progression
of lift coefficient is seen to occur monotonically decreasing
from approximately 0.94 value to approximately 0.78 at about
48 chords of flow, at which point the modified wall model was
reinstated and a monotonic progression to the final value of
approximately 0.96 is shown. No indication of unsteady
behavior was seen for this set of calculations.
The inability of the newly proposed wall model to accurately
predict the amount of separation or the increased amount of
separation for the 6.2 deg case at Mach 0.8 is disheartening.
This failure leads us to look even lower into the boundary layer
at the near wall model. Runs have not been made changing the
value of A+ because of the inappropriately long computation
-7-
I times associated with the Navier-Stokes code. Perhaps the bestarena for discussing this type of turbulence wall modification
is in the inverse boundary layer code where numerous runs can
be achieved in order to investigate parametric variations
of varying near wall modeling changes. it should be borne in
mind when discussing turbulence modeling changes such as those
above that in the thin boundary layers on airfoils, measure-
ments of the Reynolds shear stress terms will be difficult.
Measurements using today's known instrumentation technology
within the sub-layer portion of the boundary layer will be
impossible due to the large size of the instrumentation volumes
in comparison with the scales of the motions to be investigated.
This leads us to consider experiments involving extremely thick
boundary layers undergoing adverse pressure gradients in order
to allow state-of-the-art instrumentation to be brought to bear
on turbulence modeling in the near wall region. Without proper
data in the near wall region, turbulence modeling efforts will
progress very slowly. Attention should be given to the possibility
of making measurements in the boundary layers of the very large
scale facilities including, but not limited to, the NASA Ames
6 1 x6' Supersonic Wind Tunnel, the 11'x11' Transonic Tunnel, and
perhaps the 14' Transonic Tunnel.
one final calculation was made during the present reporting
period in order to add credence to the Navier-Stokes solution
}	 obtained by King using the Steger code for the case of the
i
64AO10 at M-0.72 and 3.5 deg angle of attack. This is a test
case that is being considered by King and Murphy for the
coupled TAIR and inverse boundary layer codes. The solution
shown in Figure 7 has a time variation in the upper surface
pressure distribution as shown at the various computation
times. It is also of passing note that most of these solutions
represent shock-free upper surface pressure distributions.
They are generally found to be in agreement with those
solutions given by the Steger code and in substantial disagree-
ment with those given by the coupled code.
-9-
AIRi'011. PRESSURE DISTRIBUTION
SECTION 4WO10 ALPHA` ¢O N1 p, 8_ BC
r 0.7Z	 Z^i••,t x ZS
r
r
OPJGINAL PACE 19
OF POOR QUAL rry
TURK MODEL^^
1 1...:	 _.	 t'	 ...... ' ._....:	 .....
....	
i.	
..	
_+.	 ^:
_	 ;i
	
- 1 . 0
	=	 --	
_..t
F	 s
Ld
...... 
_	 _::_
	
1.0	 1:
4	 .2	 .4	 .6	 .8	 1.0
X/C
Figure 1. Calculated pressure distribution for the NACA 64AO10
at M=0 . 8 and a-4.0 deg; Escudier model.
fs
10
vu
Z^
N
•W
OW
^4
Xj
:Y
:O
t	 +T	 F
it
	
w	 ^{	 !	 !jj
	
7	 777-
 -	 +,il fi r•	 t..;i_
	
-	 I	 I+^
	
-	
^ -
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tj
 
M 1--	 -	 ^r	 • ,-rte	 W
	
{	 t	 -;4;
_
:1- 4j
4j
	
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44
	
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R-:
}
^_ 	 tf 
1 ` fi 	 __
^.	
r+ t	 +	 I	
--	 a
r^	 y	 _.	 +
	
t	
w 
rte$ y 
H+
	
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-	
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3
rt .++	 ^. _
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rl-r 1	 t	 +.,	 t .	 r	
«+ } f+',
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-	 r	
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i-	 1	 ..	 l	 _	 ^ V
	
V	 -	 U
	
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-	
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-	 tY	 .fit	 +	 i	
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+ i 	 _	 1	 ht	 --f -	 1	 _`, I	 t	 I	 ^^.	 L	 O •-^
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-	
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-
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N
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—_	 a
t-
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4o
Rt
0. PAGE
(.	 AIRFOIL. PRESSURE DISTRIBUTION OF PQOR QUALITY
SECT I OiV_j^ffA@10 ALPHA_ ¢O _ M D. B	 BC /r	 TURB MODEL
164ftt Z 1?5-
7-	
I
	
+	 '	 -	 !	 t	 -	 !
C
is.... 
	.... .	
t:	 t:
—	
.....
	
_ ..—.
	
0	 2	 .4	 .6	 .8	 1.0
Figure 3. Calculated pressure distribution for the NACA 64AO10
at M-0.8 and a=4.0 deg; modified wall model.
i	 `
^RIGIIVA^, Me it
AIRFOIL  PRESSURE: DISTRIBUTION 
	
OF PQQR QUAL17Y
SECTION
	 D ALPHA .2	 M	 BC	 T R MOD	 tW.ii0/	
_ 6 ^ .^	 F^ _ U B MOD E L as' ,07
-	
t	
t
^::.... t^......:i^
	 .alt
t 1:	 is	 :i:i:i•..	 . _	
.....
a	 ^^^• r•
	
-	
.^ rte' .i•
	
-
	
_ r i g	 . _°: I . ^. ^: . _..i . 
	
7.	
.	 -	 =!
=.  -
	 ==:,
4	 '	
_:._	 rte_----• --___- 
t	
...
:
0.54	 . ..	 ..... ..0
0	 .z	 .4	 .6	 .8	 1.0
Figure 4. Comparison of calculated pressure distributions using
different wall models for the NACA 64AO10 at M=0.8
and a=6.2 deg.
14	 - -1
0
14
43
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0
41
44
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 7. . . . . . . . .
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=7
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. . . .	 .	 . .
77
.	 . . .
-71 -77
0.5
1.0
ORIGIM PAGE 18
	
AIRFOIL PRESSURE DISTRIBUTION
	 OF POOR QUAM
SECTION //'^,,^^^^^^ d ALPHA ,^.^ AI	 •
.G?  	 _	 . _. 77 	 I^c:^^
	
TliRE3 MODEL	 c&-
I	
-2.0 -
"7" elf	 ......f.7
t	 : ►
 :;^i.{::	 ^';	 i : ;
	 t ...-.ice .Y^iS::.^'
.... i............
	 _{.........1
.t
r
}
0	 .a
1C
Figure 6. Time variation of calculated upper surface pressure
distribution for the NACA 64A010 at M =0.72 and
a=3.5 deg, Escudier motel.


Computations and turbulent flow modeling in support of helicopter rotor technology

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