--

NASA/TM-2000-210S21 / /V / V' 1 

"'~././I /} V-:;,.2-

Investigation of Flow Separation in a 

Transonic-Fan Linear Cascade Using 

Visualization Methods 

Jan Lepicovsky 

Dynacs Engineering Company, Inc., Brook Park, Ohio 

Rodrick V. Chima, Thomas A. Jett, Timothy J. Bencic, and Kenneth E. Weiland 

Glenn Research Center, Cleveland, Ohio 

December 2000 

https://ntrs.nasa.gov/search.jsp?R=20010016660 2018-02-06T20:37:59+00:00Z
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NASA/TM-2000-210S21 

Investigation of Flow Separation in a 

Transonic-Fan Linear Cascade Using 

Visualization Methods 

Jan Lepicovsky 

Dynacs Engineering Company, Inc., Brook Park, Ohio 

Rodrick V. Chima, Thomas A. Jett, Timothy J. Bendc, and Kenneth E. Weiland 

Glenn Research Center, Cleveland, Ohio 

Prepared for the 

9th International Symposium on Flow Visualization 

sponsored by the Heriot Watt University 

Edinburgh, Scotland, United Kingdom, August 22- 25,2000 

National Aeronautics and 

Space Administration 

Glenn Research Center 

December 2000 
Acknowledgments 

The authors would like to acknowledge many useful discussions with Dr. E.R McFarland and Mr. J.R Wood and 

engineering support provided by Mr. R Torres of NASA Glenn Research Center. This program was supported 

by NASA Glenn Research Center under the Smart Green Engine program managed by Mr. RD. Corrigan. 

NASA Center for Aerospace Information 

7121 Standard Drive 

Hanover, MD 21076 

Price Code: A03 

Available from 

National Technical Information Service 

5285 Port Royal Road 

Springfield, VA 22100 

Price Code: A03 

Available electronically at http: //gltrs.grc.nasa.gov IGLTRS 
INVESTIGATION OF FLOW SEPARATION IN A TRANSONIC-FAN 

LINEAR CASCADE USING VISUALIZATION METHODS 

1. Lepicovsky 

Dynacs Engineering Compa11Y, Inc. 

Brook Park, Ohio 44142 

R.Y. Chima, T.A. lett, T.l. Bencic, and K.E. Weiland 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

ABSTRACT 

An extensive study into the nature of the separated flows on the suction side of modern 

transonic fan airfoils at high incidence is described in the paper. Suction surface flow separation 

is an important flow characteristic that may significantly contribute to stall flutter in transonic 

fans. Flutter in axial turbomachines is a highly undesirable and dangerous self-excited mode of 

blade oscillations that can result in high cycle fatigue blade failure. The study basically focused on 

two visualization techniques: surface flow visualization using dye oils, and schlieren (and 

shadowgraph) flow visualization. The following key observations were made during the study. For 

subsonic inlet flow, the flow on the suction side of the blade is separated over a large portion of the 

blade, and the separated area increases with increasing inlet Mach number. For the supersonic 

inlet flow condition, the flow is attached from the leading edge up to the point where a bow shock 

from the upper neighboring blade hits the blade surface. Low cascade solidity, for the subsonic 

inlet flow, results in an increased area of separated flow. For supersonic flow conditions, a low 

solidity results in an improvement inflow over the suction surface. Finally, computational results 

modeling the transonic cascade flow field illustrate our ability to simulate these flows numerically. 

List of symbols 

c [rnm] blade chord (89.2 rnm) 

h [rnm] blade height (95.9 rnm) 

i GM [dg] geometric flow incidence angle (10.0 dg) 

MaIN [1] inlet Mach number 

s [rnm] blade pitch (58.4 rnm) 

x [rnm] axial distance (cascade) 

y [rnm] pitchwise distance (cascade) 

r [dg] blade stagger angle (60.0 dg) 

n es [1] cascade pressure ratio 

e [dg] leading edge camber angle (-9.5 dg) 

q> [rnm] chordwise distance (airfoil) 

1 BACKGROUND 

Flutter in axial turbomachines is a highly undesirable and dangerous self-excited mode of 

blade oscillation. Modem turbine engines employ transonic fan stages with high aspect ratio blades 

that are prone to flutter and consequently suffer high cycle fatigue. Therefore, it is imperative to 

understand the origins of flutter for reliable and safe operation of these engines. Flutter can occur 

in several different operating regimes [1]. High subsonic and transonic torsional stall flutter, which 

occurs near the fan stall limit line at speeds up to about 80% of the design speed and with high 

NASAlTM- 2000-210521 
incidence, has been particularly difficult to model with analytical prediction methods. Previous 

investigations focused on instability of local supersonic regions on the blade suction surface that 

are terminated with nonnal shocks [2, 3]. It was assumed that under certain conditions the shock 

wave motion can introduce an additional contribution to the fluctuating forces on the airfoil and 

thus influence the stability of the system. Follow up investigations, however, showed that the 

shock wave dynamics associated with transonic stall flutter was insufficient to imply instability in 

the cascade based on the simple shock motion model [4]. 

The latest view of the transonic stall flutter origin is that the blade oscillations are triggered 

by high frequency changes in the extent of the partially separated area on the airfoil suction side. 

This is supported by the fact that for some of the modem tip-section airfoils at large incidence 

(about 10 dg), the shock waves do not appear until very high subsonic inlet Mach numbers have 

been reached (above 0.95). These inflow conditions are actually at the very upper end of the inlet 

Mach number range where transonic stall flutter is observed and therefore it is unlikely that flutter 

is caused by oscillating shocks. Modem transonic airfoils, however, exhibit separated flow regions 

for subsonic inlet conditions with high incidence [5, 6]. There is a lack of experimental data 

describing the separated flow characteristics of modem airfoils for transonic fans. The present 

study has been carried out to fill in this gap. 

2 TECHNICAL APPROACH 

An important unresolved question involving the flow over the suction side of a modem 

airfoil for transonic fans is the extent of the separated flow region for a subsonic inlet condition 

with high incidence. Transonic fan airfoils are designed for precompression, and therefore they 

have a sharp leading edge, a concave suction surface just downstream of the leading edge, and a 

very low overall camber. The airfoils are prone to flow separation at off-design conditions. It is 

very difficult to determine the presence and extent of the separated flow zone just from static 

pressure measurements on the airfoil surface. The separated flow region is at best recognizable 

through a plateau in the static pressure distribution that can be detected only by comparison with a 

static pressure distribution for attached flow. The extent of the flow separated zone can be 

determined reliably either by directly measuring the flow velocity distribution just above the 

surface in question or by applying one of the surface flow visualization techniques. Direct velocity 

measurement, however, requires extensive probe traversing and can be quite expensive and time 

consuming. Two visualization experimental techniques were used to detennine flow behavior on 

the suction side of an airfoil. They included: (1) surface flow visualization using dye oils, and (2) 

schlieren and shadowgraph flow visualization. Before describing the experimental techniques, the 

test facility will be introduced. 

2.1 NASA transonic flutter cascade facility 

The NASA transonic flutter cascade facility has already been described in detail in several 

publications [2 through 6]; only a brief description is given here for completeness. A schematic 

diagram of the cascade facility is given in Fig. 1. Room air enters the bell mouth (top left), passes 

through the test section (inserted photograph), and exits through a diffuser (bottom right) into a low 

pressure exhaust system. The facility operates in a continuous mode. The test section consists of 

nine blades. The setting angle of the inlet duct is 20.0 dg, and the blade-setting angle is 30.0 dg; 

this results in 10.0 dg chordal incidence for the airfoils. The exit channel is set at an angle of 24.0 

dg. This configuration assures a nearly identical flow periodicity for six blades in the cascade 

(from blade B2 to blade B7; counted from left to right). A complete analysis of the cascade steady 

state flow periodicity is given in references 6 and 7. Airfoil geometry and cascade geometry is 

given in Fig. 2 [5, 6]. Flow rate through the cascade (inlet Mach number) is controlled by valves 

located downstream of the exit diffuser. All the experiments described in this paper were carried 

out for steady state conditions. It should also be stated here that no boundary layer bleed was 

employed during this study. 

NASNTM-2000-21 0521 2 
· f ~:.r-

" . / /' ....... . "'; j";'T 

,,'- , I . . 

I INLET WALL 

\. ANGLE 20 d91-".·.., ":<:;;;;;;i:;;;,o;;;;;';;;;~;;.:;,;,;,;~~;;;; 

I ' ./ ••,' ' BLAD . E SETTING----- ~ . 

- .. '-~< ANGLE 30 dg-----___ ~ 

.. , ... -I' I '. EXIT W;~~-----__ EXIT 

V'. . - ,.'"., .. " { \.. II i'.'i'.' ANGLE 24 dg -----_ .. oj 

"., i 

r: 

V. 

Fig. 1. NASA GRC transonic flutter cascade 

2.2 Dye-oil surface flow visualization 

PITC>fNISE QlRECTlON 

+y 

Fig. 2. Airfoil and cascade 

cool'dinate system 

Surface flow visualization teclmiques are based on a dye, which is of a liquid or a solid 

form, being smeared over the surface by flow. Dye dots or lines are made on the test surface just 

before the test facility is started. This is relatively simple and straightforward. However, very 

often, it is difficult to distinguish which of the resulting dye traces were made during the facility 

start-up process and which were made while running at the test conditions. The shut-down 

transient is usually no problem because the facility can run until the dye traces are sufficiently dry. 

To avoid the start-up confusion, some researchers add solid particles in the flow after the test 

conditions have been established and rely on particle deposition on the test surface [8]. The 

heaviest depositions are usually along the line of flow reattachment or in the regions of high 

turbulence intensity. The other approach is to make marks on the test surface with an easily soluble 

color and then to spray a solvent in the flow at desired operating conditions. Both approaches were 

tried but without success and therefore an old-fashioned, reliable dye-oil technique was employed. 

Using the dye-oil technique, droplets of colored oil are deposited on the test surface and the 

facility is started. Very often, particularly for low viscosity oils and a slow facility start-up, the 

recorded flow picture captures the transient flow patterns that occur .during the tunnel start-up 

phase rather than the test flow condition pattern. To avoid this problem, two aspects must be 

resolved: (1 ) very short facility start-up time, and (2) a dye-oil mixture of suitable viscosity (in 

conjunction with small dye-oil marks on the test surface). To minimize the start-up times, the 

cascade facility operating conditions were set up using the control valves as usuaL Then, while the 

control valves stayed open, the facility was turned on and off using the main shut-off valve only. 

Through this procedure, the maximum Mach number of 1.2 was reached within 6 s from the time 

the main shut-off valve started to open. This extremely short start-up time caused vibrations in the 

facility and should not be overused; however, so far no detrimental effects on the facility have been 

noticed. Commercial oil-based paints were mixed with automotive oil viscosity SAE 85W140. 

The oiVpaint mixture ratio was between 3:2 and 2: 1, depending on the paint color. The right 

mixture ratio was determined for each color by testing the color mark fluidity. Color dots 2 to 3 

mm in diameter were made on a test surface, and the surface tilted in the vertical direction. Oil was 

added to the mixture until the color marks started to run after the test surface was tilted. 

In preparation for tests, color dots were made on a test blade mounted in the cascade. The 

blade-setting angle was 30.0 dg with respect to the horizontal direction. For the high solidity 

cascade, the spacing between the blades is quite smalL The dots were made using a l-mmdiameter copper wire that was dipped in the mixture. The color marks on the cascade sidewall 

were made on a segment that was removed from the wall as shown in Fig. 3. Behavior of the color 

dots during the facility start-up procedure was observed to verify that the recorded flow pattern is 

not contaminated by the transient phenomenon. As mentioned above the flow in the cascade 

reached the test condition in 6 s. During this period, and for an additional 10 to 15 s, the fluid in 

the dots was pushed to the downstream side of the dot. Approximately 20 to 25 s after flow onset 

the color dots started to break and thin streaks ran over the blade surface. It took about 3 minutes 

NASN TM- 2000-21 0521 3 
for all the fluid from the dot to smear onto the blade surface. The flow was maintained at the test 

condition for 15 mllutes in order to dry the paint. Only one color or one line of dots was made at a 

time on the blade tested. Each of the multicolor surface flow pictures to be discussed below was 

completed in 5 or 6 consecutive runs. The consistency of the flow pattern for different colors 

proves that the flow in the cascade is very stable and repeatable. Also, it proves that the recorded 

pictures are of the flow at the steady state conditions and that the pattern is not affected by any 

tunnel start-up flow transients. After the test was completed, the blade was dismounted from the 

cascade and photographed to record the surface flow pattern. 

2.3 Flowfield visualization 

LIGHT SOURCE 

WINDOW 

SPHERICAL 

MIRROR 

Fig. 4. Double-pass schlieren system 

Fig. 3. Color dots being made on a test article 

The double-pass schlieren system used in thjs study is depicted in Fig. 4 [2, 4]. The blades 

are mounted between end walls with a mllror on one side and optical quality glass window on the 

other side. The mirror and the window comprise a part of the schlieren system. The limitation in 

the region of flow visualization over the airfoil surface is due to the large blade trunillon and the 

blade end fillets. Difficulties were experienced with the alignment of the schlieren system. A 

perfect alignment of the knife-edge before the test would rapidly deteriorate once the flow started. 

It is believed that thjs was caused by mllute deflections of the facility walls due to the pressure 

difference across the tunnel walls. Also, there was a noticeable unsteadiness in the image position 

at the knife-edge location. For this reason, most of the pictures with shock wave structures present 

were acquired as shadowgraphs. The flow images were recorded either on a 35-mm camera or a 

CCD can1era with resolution of 512 by 512 pixels. 

3 SURFACE FLOW VISUALIZATION RESULTS 

Solidity is an important parameter in cascade design. It is defined as a ratio of blade chord 

and pitch. In turbomachinery, high solidity is used to reduce loading and confine the flow, which 

reduces separation. However, high solidity increases wetted area and losses in a cascade, therefore 

there is an optin1Um for solidity of each cascade. The effects of solidity on surface flow pattern are 

discussed next. 

3.1 High solidity cascade 

The effects of the inlet Mach number on the evolution of the surface flow pattern on the 

airfoil suction side are shown in Fig. 5. The results are for the high solidity cascade of cis = 1.52. 

For subsonic Mach numbers of 0.5 and 0.8, the flow exhibits a large separated region that starts 

immediately at the leading edge and extends, at midspan, down to 47% and 52% of the blade 

chord, respectively. A schematic diagram explaining the flow pattern for subsonic conditions is 

shown in Fig. 6. For inlet flow at the low supersonic Mach number of 1.18 (Fig. 5), the surface 

flow pattern is completely different. First, starting from the leading edge, the flow is attached to 

the blade surface down to approximately 32% of the chord. Then, there is a separated flow region, 

NASAlTM-2000-21 0521 4 
MaiN = 0.50 

MaiN = 0.80 

INLET 

FLOW 

INLET FLOW 

MaiN = 1.00 

HIGH SOLIDITY 

CASCADE 

cIs = 1.52 

MaiN = 1. 18 

Fig. 5. Effects of inlet Mach number on suction surface flow pattern for high solidity cascade 

in which air moves in the direction against the inlet flow. Finally, at a locati on 62% of the blade 

chord, the flow attaches back to the blade surface . The fl ow behavior can be understood better in 

Fig. 7, where there is a combination of shadowgraph and surface fl ow visualization pictures. 

Starting with the shadowgraph picture, three types of shock waves can be detected here. First, 

there is a lip (oblique) shock, attached to the blade leading edge. This shock is a consequence of 

flow overexpansion around the leading edge and adj usts the flow direction along the blade suction 

surface. As will be shown later, this shock depends solely on the flow conditions at the blade 

leading edge. The second shock, the bow shock, in particular its location, depends strongly on the 

inlet Mach number and the cascade solidity. Finally, the trailing edge shock depends on the 

pressure ratio across the cascade. The oversized blade end fill et blocks the view of the bow shock 

at the blade suction surface . The proj ected impact point of the bow shock on the blade surface is 

indicated by the red-line extension of the shock (the fact that the shock most likely splits into a 

lambda pattern is ignored here for simplicity). The location of the bow shock impact is redrawn in 

scale on to the surface flow pattern from the shadowgraph picture. The excellent agreement with 

the line of flow separation is clearly demonstrated here. The flow separates slightly upstream of 

the shock due to the secondary boundary layer fl ow ' under the shock' . The position of the 

reattachment line (at 62% of the blade chord) is estimated from the surface flow picture. 

3.1.1 Subsonic inlet Mach numbers 

The effects of cascade so lidity on the extent of flow separation were investigated for a high 

subsonic in let Mach number of about 0.8. The results are shown in Fig. 8. For the high so lidity 

cascade, the fl ow on the airfoil suction side is separated at the midspan up to approximately 52% of 

the airfoil chord. The separated region is restricted to a parabola, whi ch is symmetric along the 

blade midspan line and extends to the blade upstream comers. Two swirling traces fo rmed 

downstream of the blade leading edge comers and rotating in opposite directions are clearly visible 

here. 

The cascade so lidity was lowered to 0.76 by removing the even number blades (the cascade 

was now composed of fi ve airfo ils only), and the experiment was repeated. On comparing the 

results with the previous case, it is obvious that fo r the lower cascade solidity the separated region 

is larger and extends down to approximate ly 65% chord at the blade midspan. Furthermore, even 

ASAlTM- 2000-21052 1 5 

i -
SUBSONIC ~ INLET 

FLOW 

AT HIGH 

SEPARATED FLOW ZONE 

SECONDARY A TTACHED SECONDARY 

FLOW ZONE FLOW FLOW ZONE 

Fig. 6. Surface flow pattern on blade 

suction side for subsonic flow 

though the boundary I ine of the entire separated region 

seems to be symmetric along the blade midspan line, a 

closer look reveals that there is an asymmetry in flow 

traces on the blade. There is an additional swirling 

pattern in the upper right region that rotates counter 

clockwise (left of the comer swirling trace that rotates 

in the clockwise sense). Clearly, there is no symmetry 

of the fl ow pattern within the separated fl ow region. 

Finally, the experiment at a high subsoni c inlet 

Mach number was carried out for a single airfo il (blade 

85 only). The separated flow region now covers most 

of the blade suction surface and extends down to the 

airfo il trailing edge. Only the downstream blade 

corners show small regions of attached flow. There are 

again only two counter-rotating swirling traces in the 

upstream (leading edge) corners, similar to the high 

solidity cascade case. The boundary of the separated 

BLADE B5 TOPVIEW ___ _ 

CASCADE SIDEVIEW 

INLET 

FLOW LIP 

INLET 

FLOW 

> HO K /BOW SHOCK 

BOW SHOCK 

@ O.32 * C 

REA TTACHMENT 

LINE @ O. 62 * C 

Fig. 7. Surface flow pattern for supersonic inlet flow, MaIN = 1. L8 

region, however, is not symmetric along the blade mid span line. The separated flow region IS 

skewed to the left (in the sense of the in flow direction). 

3. 1.2 Supersonic inlet Mach numbers 

The results fo r the supersonic inlet Mach numbers are depicted in Fig. 9. As already 

indicated in Fig. 5, the surface flow pattern fo r supersonic inlet flow fo r the high solidity cascade is 

dramatically diffe rent from the subsonic inlet conditions. The features of the flow patte rn for 

supersonic inlet and the high solidity cascade were discussed in detail in Fig. 7. For the low 

so lidity cascade (see Fig. 9) the flow pattern is essentially similar to the high so lidity case with the 

diffe rence that now the bow shock impinges on the blade surface very close to the trailing edge . 

The traces of flow separation caused by the shock wave / boundary laye r interaction are visible at 

about 93% of the blade chord. It should be noted that the shock impingement on the blade is 

visibl e only on l8% of the blade span (in comparison 56% for the high so lidity cascade). The flow 

is attached on most of the blade suction surface, with secondary fl ow zones visibly migrating close 

to the blade midspan line. 

The last flow pattern in Fig. 9 is for the case of a single airfo il. Clearly the flow is attached 

over the entire blade surface . There are no traces of flow separation visible. The secondary fl ow 

regions migrate toward the blade center in a similar manner to the two previous cases. While in the 

NASN TM- 2000-2 10521 6 
previous two cases the flow pattern was symmetric along the blade midspan line, for the single 

airfoil the fl ow pattern is skewed to the left (in the flow direction), as it was for the case of the 

subsonic inlet flow. 

INLET FLOW 

HIGH SOLIDITY 

cIs = 1.52 

MaiN = 0.80 

INLET FLOW 

LOW SOLIDITY 

cIs = 0.76 

MaIN = 0.81 

INLET FLOW 

SINGLE 

AIRFOIL 

MaiN = 0.85 

Fig. 8. Effects of cascade solidity for subsonic inlet fl ow conditions 

HIGH SOLIDITY 

cIs = 1.52 

MaiN = 1. 18 

INLET 

FLOW LOW SOLIDITY 

cIs = 0.76 

MaIN = 1.26 

INLET 

FLOW SINGLE 

AIRFOIL 

MaiN = 1.28 

Fig. 9. Effects of cascade solidity for supersonic inlet fl ow conditions 

4 SHOCK WAVE STRUCTURE 

The goal of the flow visualisation experiment was to determine the onset of shocks in the 

flow for an increasing in let flow Mach number and to observe the changes in shock wave structure. 

The experiments were carried out for a range of in let Mach numbers starting at 0.6. Three basic 

configurations were tested: (I) high so lidity cascade, cis = 1.52; (2) low solidity cascade, cis = 0.76; 

and (3) a single airfoil. 

4.1 High solidity cascade 

The sequence of photographs in Fig. 10 shows the results for the case of the hjgh solidity 

cascade. For the subson ic Mach numbers of 0.6 and 0.9 (Fig. 10a,b) there are no shocks present in 

the flow. Obviously, the critical Mach number for this cascade is above Mach number 0.9. Local 

regions of high and low pressure in the vicinity of the blade leading edges are visible here (light 

AS TM- 2000-21052 1 7 
and dark areas). Also, the regions of separated flow on the suction surface, just downstream of the 

leading edges, can be detected. Wakes downstream of the blades are clearly visible. 

The first appearance of shock waves in the flow occurs for the inlet flow Mach number of 

1.0 I (Fig. 10c). As seen here, the shock structure is not periodic; each blade shows a different 

MaIN = 0.60 1Ccs = 1.036 

MaIN = 0.90 1Ccs = 1.089 

MaIN = 1.01 1Ccs = 1.116 

MaIN = 1.05 

MaIN = 1.12 1Ccs = 1.120 

MaIN = 1.21 7rcs = 0.869 

g. J 

Cascade solidity cis = 1.52 

Incidence I GM = +10.0 dg 

Fig. 10. High solidity cascade shock 

patterns 

shock pattern. Blades B4 and B5 generate simple normal shocks, whereas on blades B6 and B7 the 

shock structure appears to consist of an oblique shock fo llowed by a bow shock. The location of 

the bow shock, particularly on blades B6 and B7 varies significantly. Direct obse rvation of the 

shock structure for this inlet Mach number revealed that the shock structure was highly unstable 

and varied rapidly. Once the inlet Mach number was raised to 1.05, the shock structure noticeably 

stabilized and exh ibited the pattern shown in Fig. 10d. The shock pattern appears to be periodic 

with a period equal to two blade pitches. Even blades (B4 and B6) generate normal shocks at 40% 

of the blade chord, whereas odd blades (B5 and B7) clearly show a horizontally located oblique 

(lip) shock attached to the blade leading edge and a normal shock (bow) at 25% of the blade chord. 

ASAlTM- 2000-21 0521 8 
_. -- ~~~~ -~---

For the inlet Mach numbers of l.l2 and higher (Fig. IOe,f,g) the shock structure is highly periodic 

with the period of one blade pitch. Additional increase of the inlet Mach number to 1.17 (Fig. lOt) 

slightly shifts the bow shock in the flow direction. Also, the slip lines separating two subsonic 

flow streams of different velocity and entropy become clearly visible. Finally, at this operating 

condition one can observe the fomlation of trailing edge shocks on blades B5 and B6. At the last 

test condition (Mach number 1.21 , Fig. 109) the flow in the cascade is choked and the cascade does 

not generate any pressure increase. The shock wave pattern at the cascade inlet does not change. 

The shocks in the trailing edge region become more intense and extend more deeply into the 

flowfield. 

4.2 Low solidity cascade 

The evolution of the shock structure for the low solidity cascade with increasing inlet Mach 

number (Fig. 11) exhibits some features that are similar to the high solidity case. The shocks 

appear for the fust time in the flow for an inlet Mach number of 0.97 (Fig. Ila). Blade B5 exhibits 

a bow shock whereas blade B7 still seems to be shock-free. Separated flow is traceable on both 

blades just downstream of the leading edges. At the inlet Mach number of 1.02 (Fig. 11 b), a 

combination of a lip and a bow shock is generated on blade B5. There is no flow separation at the 

blade leading edge of blade B5. Blade B7 shows traces of unstable shocks above the suction 

surface and the flow is visibly still separated at the blade leading edge. A gradual increase of the 

inlet Mach number (Figs. 1 Ic,d,e) stabilizes the shock pattern in the cascade. Both blades show a 

combination of lip and bow shocks. On all of these figures, the shock system on blade B7 always 

appears weaker, as though blade B7 was facing a lower inlet Mach number than blade B5. An 

interesting shock formation can be detected for the inlet Mach number of 1.19 (Fig. 11 t) and higher 

(Figs. 11g,h). First, a trailing edge shock appears on the pressure side of the airfoil, and with an 

increasing inlet Mach number this trailing shock penetrates more deeply into the flowfield. 

Second, and more interestingly, there seems to be an additional oblique shock generated at 

midchord on the suction side. This shock interacts with the lip shock for inlet Mach numbers of 

1.19 and higher. The interaction is manifested by a sudden direction change of the lip shock line. 

4.3 Single airfoil 

The last sequence of pictures presented in Fig. 12 shows the flowfield evolution for a 

single airfoil. Flow is separated over the airfoil suction side at least up to the inlet Mach number of 

0.97 (Fig. 12a). The transition to the lip-shock / bow-shock combination seems to be quite rapid; 

for the inlet Mach number of 1.04 (Fig. 12b) the combination is already in place. Also, the trailing 

edge shock waves appear at a lower inlet Mach number than was the case for the low and high 

solidity cascades; first on the airfoil pressure side at a Mach number of 1.09 (Fig. 12c), and then 

also on the suction side starting at a Mach number of 1.13 (Fig. 12d,e,t). As the inlet Mach number 

is raised, the suction side bow shock moves in the downstream direction toward the trailing edge 

(Fig. 12g). There is a visible interaction of the lip shock with an oblique shock from the blade 

suction surface starting at the inlet Mach number of 1.20 (Fig. 12t). This is similar to the case of 

the low solidity cascade as shown in Fig. 11g,h. For the highest inlet Mach numbers (1.26 and 

higher), the only change in the shock wave structure is a decreasing angle of the trailing edge shock 

waves as the inlet Mach number is raised. 

5 COMPUTATIONAL RESULTS 

A numerical investigation was also carried out in parallel with the experimental effort. 

Even though the computational methods and results are not the topic of this paper, the following 

illustrates some of our capabilities and accomplishments. 

Two computational fluid dynamics codes have been used to analyze the blades described 

here. The fust code, RVCQ3D [9] analyzes viscous flow around an isolated blade on a twodimensional blade-to-blade plane. The second code, Swift [10] analyzes fully three-dimensional 

blades using a multiblock grid. Both codes solve the Navier-Stokes equations using an explicit 

finite-difference scheme and an algebraic turbulence model. 

NASA/TM-2000-21 0521 9 
-------

Ma'N = 0.97 1CGS = 1.155 Ma'N =1 .16 1CGS = 1.152 

Ma'N = 1.02 1CGS = 1.146 Ma'N = 1.19 1CcS = 1.144 

.• '(. 

b. 

~ c. 

1CGS = 1.159 

d. 

Fig. 11. Low solidity cascade shock patterns 

Figure 13 shows a solution computed by RVCQ3D for an inlet Mach number of 1.2 and an 

incidence angle of 9.0 dg. The figure shows contours of the magnitude of the density gradient, 

which may be compared to the experimental shadowgraph photo shown in Fig. 109. A strong bow 

shock wave stands far ahead of the leading edge and hits the neighboring blade normal to the 

suction surface, separating the boundary layer. A weak lip shock is generated at the leading edge 

where the blade curvature changes from round to concave on the suction surface. A weak slip line 

is barely visible where the two shocks intersect at mid passage. Figure 14 shows a solution 

computed using Swift for a Mach number of 1.10 and an incidence angle of 10.0 dg. The figure 

shows particle traces just above the suction surface, which may be compared to the experimental 

surface oil flow patterns shown in Fig. 5 (MaIN = 1.18). The flow is from top to bottom. The 

supersonic inflow expands around the leading edge and remains attached (red.) The flow separates 

at the shock (red-blue) and reattaches near mid chord (blue-white.) Secondary flows between the 

separation bubble and sidewall fIllets are also seen (white). 

NASA/TM- 2000-21 0521 10 

l 
MaIN = 0.97 Jies = 1.155 MaIN = 1.20 Jies = 1.199 

a. 

MaIN = 1.04 Jies = 1.186 MaIN = 1.24 Jies = 1.144 

b. g. 

MaiN = 1.09 Jies = 1.201 MaIN = 1.26 Jies = 1.099 

c. h. 

MaIN = 1.13 Jies = 1.211 MaIN = 1.29 Jies = 0.946 

• 

d. 

MaIN = 1.17 MaIN = 1.30 Jies = 0.892 

Fig. 12. Single airfoil shock patterns 

ASNTM- 2000-210521 II 
Fig. 13. Computed density gradient pattern 

for the inlet Mach number of 1.2 

and incidence of 9.0 dg 

Fig. 14. Computed surface flow 

pattern for the inlet flow 

Mach number of 1.1 and 

incidence of 10.0 dg 

6 CONCLUSIONS 

Several key observations were made during the course of this study. They can be 

summarized as fo llows. 

• For the subsonic inlet flow at the incidence of 10.0 dg, the flow on the suction side of the blade 

is separated over a large portion of the blade surface. The size of the separated area increases 

with the increasing subsonic inlet Mach number. For the supersonic inlet Mach number the 

flow is attached from the leading edge up to the point where a bow shock hits the blade surface. 

The bow shock interaction with the blade boundary layer results in a relative ly short separati on 

bubble. For the sonic inlet Mach number, the surface fl ow pattern on the front half of the 

suction surface exhibits a high degree of randomness that indicates instability of the local 

shocks generated in this area. 

• The effects of cascade so lidity are diffe rent fo r subsonic and supersoni c inlet conditions. A 

low cascade so lidity, fo r the subsonic inlet flow, results in an increased area of separated flow. 

For a single airfoil, the flow is separated over 80% of the suction surface. At midspan, the 

separation extends up to the airfoil trailing edge. For supersonic fl ow conditions, low so lidity 

results in an improvement in flow over the suction side because the bow shock moves in the 

downstream direction, and consequently the separation bubble caused by the bow shock 

impingement is reduced in size. For the iso lated airfo il , the flow is attached over the entire 

suction surface. 

• Local shocks over the suction surface of the high solidity cascade do not appear until a 

re lative ly high subsonic inlet flow Mach number of 0.95. The shock structure fo r the range of 

inlet Mach numbers from 0.95 to L. 05 is very unstable and not periodic. Above a Mach 

number of L.05 , the shock structure is stable and periodic. Generally, the shock structure 

consists of a lip (oblique) shock and a bow (termination) shock. It appears that the lip shock 

angle, once the shock is set, is independent of the inlet Mach number (at least fo r the range 

tested). The bow shock moves in the downstream direction as the inlet Mach number increases 

(that is, as the back pressure drops). Trailing edge shocks appear for inlet Mach numbers 

higher than I. 1 5. Similar observations were made a lso for the cases of low cascade so lidity 

and a single airfoil. 

• Computational results showed a good agreement with experimental results for the shock 

structure in the flow fo r the high solidity cascade, particularly on the upstream face of the 

cascade. The trailing edge shocks were not captured by the calculations. The predi ction of 

flow separation fo r the supersonic inlet conditions agrees qualitative ly with experiment. It 

appears that the separation bubble in the calculations starts closer to the leading edge than is 

seen in the experimental data. 

NASN TM- 2000-2 1 0521 12 
REFERENCES 

[1] Carta, F.O.: Aeroelasticity and Unsteady Aerodynamics, Aircraft Propulsion System Technology and Design, edited 

by G.C. Oates, AlAA Education Series, Chapter 7, 1989. 

[2] Boldman, D.R., Buggele, A.E., and Michalson, G.M.: Stall Flutter Experiment in a Transonic Oscillating Linear 

Cascade, NASA TM-82655, 198! 

[3] Goldstein, M.E., Braun, W., and Adamczyk, J.1. : Unsteady Flow in a Supersonic Cascade with Strong In-passage 

Shocks, Journal of Fluid Mechanics, Vol. 83, pp. 569-604, 1977. 

[4] Shaw, L.M., Boldman, D.R, and Buggele, A.E.: Unsteady Pressure Measurements on a Biconvex Airfoil in a 

Transonic Oscillating Cascade, NASA TM-86914, 1985. 

[5] Buffum, D.H., Capece, V.R., King, A.1., and El-Aini, Y.M.: Oscillating Cascade Aerodynamics at Large Mean 

Incidence, NASA TM-1 07247, 1996. 

[6] Lepicovsky, J., McFarland, E.R, Chima, RV., and Wood, J.R.: On Flow Periodicity in the NASA transonic Flutter 

Cascade, Part I - Experimental Snldy, NASA TM-2000-209934,2000. 

[7] Chima, R.V., McFarland, E.R., Wood, J.R., and Lepicovsky, J.,: On Flow Periodicity in the NASA transonic Flutter 

Cascade, Part I - Numerical Study, NASA TM-2000-209933, 2000. 

[8] Bradshaw, P.: Experimental Fluid Mechanics, Pergamon Press Ltd., 1970. 

[9] Chima, R V.: Explicit Multigrid Algorithm for Quasi-Three-Dimensional Viscous Flows in Turbomachinery, AIAA 

1. Propulsion and Power, Vol. 3, pp.397-405, 1987. 

[10] Chima, R V. : Calculation of Tip Clearance Effects in A Transonic Compressor Rotor, ASME Journal of 

Turbomachinery, Vol. 120,pp 131-140, 1998. 

NASNTM- 2000-21 0521 13 
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December 2000 TechrU ca ~emoran um 

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS 

Investigation of Flow Separation in a Transonic-Fan Linear 

Cascade Using Visualization ~ethods 

6. AUTHOR(S) VVU-523-26-13-00 

Jan Lepicovsi.)" Rodrick V. Chima, Thomas A. Jett, Timothy J. Bencic, 

and Kenneth E. Weiland 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION 

REPORT NUMBER 

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field E- 12488 

Cleveland, Ohio 44135-3191 

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AGENCY REPORT NUMBER 

ational Aeronautics and Space Administration 

Washington, DC 20546-0001 NASA TM-2000-21OS21 

11. SUPPLEMENTARY NOTES 

Prepared for the 9th International Symposium on Flow Visualization sponsored by the Heriot Watt University, Edinburgh, 

Scotland, United Kingdom, August 22- 25, 2000. Jan Lepicovsky, Dynacs Engineering Company, Inc., 2001 Aerospace 

Parkway, Brook Park, Ohio 44142 (work funded by NASA Contract NAS3- 98008); Rodrick V. Chima, Thomas A. Jett, 

Timothy J. Bencic, and Kenneth E. Weiland, NASA Glenn Research Center. Responsible person, Jan Lepicovsky, 

organization code 5810, 216-977-1402. 

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE 

Unclassified -Unlimited 

Subject Categories: 02 and 07 Distribution: Nonstandard 

Available electronically at http://gltrs.grc.nasa.gov/GLTRS 

This publication is available from the NASA Center for AeroSpace Information, 301-621-0390. 

13. ABSTRACT (Maximum 200 words) 

An extensive study into the nature of the separated flows on the suction side of modem transonic fan airfoils at high 

incidence is described in the paper. Suction surface.flow separation is an important flow characteristic that may significantly contribute to stall flutter in transonic fans. Flutter in axial turbomachines is a highly undesirable and dangerous 

self-excited mode of blade oscillations that can result in high cycle fatigue blade failure. The study basically focused on 

two visualization techrUques: surface flow visualization using dye oils, and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction side 

of the blade is separated over a large portion of the blade, and the separated area increases with increasing inlet Mach 

number. For the supersonic inlet flow condition, the flow is attached from the leading edge up to the point where a bow 

shock from the upper neighboring blade hits the blade surface. Low cascade solidity, for the subsonic inlet flow, results 

in an increased area of separated flow. For supersonic flow conditions, a low solidity results in an improvement in flow 

over the suction surface. Finally, computational results modeling the transonic cascade flowfield illustrate our ability to 

simulate these flows numerically. 

14. SUBJECT TERMS 15. NUMBER OF PAGES 

19 

Transonic airfoil; Flow separation; Surface flow visualization; Shock wave structure 16. PRICE CODE 

A03 

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