Flowfield Measurements in a Slot-Bled Oblique Shock-Wave and by liuhongmei


									NASA/TM---1998-206974                                                                               AIAA-95-0032

Flowfield                      Measurements                                    in a Slot-Bled
Oblique Shock-Wave    and Turbulent
Boundary-Layer  Interaction

D.O. Davis, B.P. Willis,             and W.R.      _gst
Lewis Research   Center,              Cleveland,      Ohio

Prepared       for the
33rd    Aerospace         Sciences       Meeting        and    Exhibit
sponsored        by the American            Institute         of Aeronautics   and   Astronautics
Reno,      Nevada,       January     9-12,1995

National       Aeronautics         and
Space      Administration

Lewis      Research       Center

April        1998

        The authors   would   like to gratefully    acknowledge          Mr. T. Bencic    of NASA      Lewis     Research     Center      for
                               matters    cor, cerning     the pressure-sensitive        paint   technique.

                                                             Available    from

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                                                FLOWFIELD    MEASUREMENTS   IN A SLOT-BLED                                              OBLIQUE
                                            SHOCK-WAVE    AND TURBULENT   BOUNDARY-LAYER                                               INTERACTION

                                                                      D. O. Davis*, B. P. Willis*, and W. R. Hingst*
                                                                      NASA Lewis Research Center, Cleveland,   Ohio

                                                     Abstract                                               ,y           =     ratio of specific         heats
                                                                                                            6            =     boundary-layer           thickness
                      An experimental             investigation was conducted  to de-                       6*           =     displacement   thickness
                termine the flowfield             inside a bleed slot used to control
                                                                                                            0            =     momentum     thickness
                an oblique shock-wave        and turbulent boundary-layer     in-
                                                                                                            P            =     density
                teraction.   The slot was oriented normal to the primary
                flow direction and had a width of 1.0 cm (primary flow
                direction), a length of 2.54 cm, and spanned 16.5 cm. The
                approach boundary layer upstream of the interaction was                                                                     Subscripts
                nominally    3.0 cm thick. Two operating conditions         were
                studied:    M=1.98 with a shock generator         deflection an-
                                                                                                            0           =      condition      in wind-tunnel        plenum
                gle of 60 and M=2,46 with a shock generator deflection
                                                                                                             1          =      condition      in Zone 1
                angle of 80 . Measurements          include surface and flow-
                                                                                                            2           =      condition      in Zone 2
                field static pressure, Pitot pressure,     and total mass-flow
                                                                                                            3           =      condition      in Zone      3
                through the slot. The results show that despite an initially
                two-dimensional     interaction for the zero bleed-flow case,                               cl          =      centerline     condition
                the slot does not remove mass uniformly       in the spanwise                               i           =      inviscid     condition
                direction.   Inside the slot, the flow is characterized     by                              ref         =      condition      at upstream        reference   plane
                two separation    regions which significantly   reduce the ef-                              W           =      condition      at wall
                fective flow area. The upper separation region acts as an
                aerodynamic     throat resulting in supersonic flow through
                much of the slot.

                                                                                                                 REVENTION        or control of boundary-layer      separa-
                                                                                                                 tion resulting from a shock-wave     and boundary-layer
                  A                                                                                        interaction can be achieved by removing low momentum
                               :      area
                                                                                                           fluid (boundary-layer      mass) near the surface    through a
                  D            =      slot width
                                                                                                           porous region. 1, 2 The porous region may be a series of
                  Hint         =      incompressible              shape factor
                                                                                                           discrete perforations    (holes) or a single slot. Numerical
                  L            =      slot length
                                                                                                           analysis    of the interaction        region        can be very    costly   in
                  rh           =      mass-flow rate
                                                                                                           terms of computer time and grid generation         if the flow
                  _n'bt        =      unit mass-flow rate in boundary-layer                                through the individual     holes or the slot is endeavored.
                  m            =      normalized   mass-flow rate                                          This is especially true if the interaction   is part of a full
                  M            =      Mach number                                                          supersonic inlet calculation.   To avoid this, a global bleed
                  P             =     static      pressure                                                 model is sought which eliminates     the need for resolving
                  Pt            =     total pressure                                                       the bleed flow passages.  An effective global bleed model
                  P_2           =     Pitot pressure                                                       should do two things:    predict the amount     of mass re-
                                                                                                           moved from the boundary-layer     and predict the condition
                  Pw            =.     normalized  wall static              pressure
                  Q.            =      sonic flow coefficient                                              of the boundary-layer    downstream    of the interaction. The
                                                                                                           former is important from the standpoint          0f bleed drag
                  Re            =      Reynolds        number
                                                                                                           and bleed system scaling while the latter is an indica-
                  S             =      slot span
                                                                                                           tion of the effectiveness    of the bleed on maintaining      a
                  Tt            =      total temperature                                                   "healthy" boundary-layer.     To achieve these goals, the lo-
                  x,y,z         =      cartesian      coordinate system
                                                                                                           cal flow phenomena         in the interaction   region including
                                =      distance      from nozzle exit to slot I.e.
                  Xslot                                                                                    the bleed passages and bleed plenum must be understood
                   Ol           =      shock      generator        deflection     angle                    and accounted for in a model. Not surprisingly,         previous
                                                                                                           investigations   resolving    the flow through bleed passages
                                                                                                           have been computational          in nature.   Hamed et al. have
                       *Research Engineer,               Inlet,     Duct,       and Nozzle   Flow          studied the oblique shock-wave           and laminar boundary-
                 Physics Branch.                                                                           layer interaction with _  bleed through normal slots, 3 and the

                 NASA/TM--1998-206974                                                                  1
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                                                  Fig. 1 Schematic                                                              of bleed                                                                                                   Fig. 3 Plan view of slot bleed experiment.
                                                 interaction region                                                           (no scale).
                                                                                                                                                                                                                           region(s) in the bleed passage which have a large effect
                                                                                                                                                                                                                           on the flow coefficient of the hole or slot 13 and the :pres-
                     oblique shock-wave and turbulent boundary-layer   interac-
                     tion with bleed through various normal 4-6 and slanted 5, 7                                                                                                                                           ence of a two-segment       b_ier     shock (see Fig,: i) which
                                                                                                                                                                                                                           may or may not :be_ attached         depending     on local flow
                     slots. Because the slot configuration   is two-dimensional,
                                                                                                                                                                                                                           conditions,   Although these studies have provided a sig-
                     Hamed et al. were able to perform a fairly comprehen-
                                                                                                                                                                                                                           nificant increase, in the understanding       of the shock:wave
                     sive paramemc    study including the effects of bleed mass-
                                                                                                                                                                                                                           and boundary-layer     interactionwith      bleed, unfortunately
                     flow rates, slot location relative to shock impingement
                     location, slot inclination  angle, and slot length-to-width                                                                                                                                           there has been no experimental       flowfield data with which
                     ratio. Hahn et al. 8 also numerically   studied the oblique                                                                                                                                           to compare. The present experimental           investigation  was
                                                                                                                                                                                                                           under_en      in order to gain further insight into the slot.
                     shock-wave and turbulent boundary-layer    interaction with
                                                                                                                                                                                                                           bled oblique shock-wave       and turbulent boundary-layer      in-
                     bleed through normal slots and in addition to investigating
                                                                                                                                                                                                                           teraction and to provide validation data for CFD methods.
                     various shock impingement    locations and slot length-to-
                     width ratios included the effects of two slots with vari-
                     ous streamwise   spacing.   Rimlinger,  Shih and Chyu 9-12                                                                                                                                                                                            Experimental                                                          Program
                     considered the more complex (three-dimensional)      case of
                     oblique shock-wave and turbulent boundary-layer     interac-                                                                                                                                                A schematic of the slot bleed experiment                                                                                                                 with ref-
                     tions with bleed through holes. _ese      numerical  studies                                                                                                                                          erence coordinates is shown in Figs. 2 and 3.                                                                                                                 The bleed
                     have identified several important features of the bleed in-                                                                                                                                           model is a single 900 slot, 1.00 cm wide (D)                                                                                                                   and 2.54
                     teraction.   Among then-, are the presence    of separation                                                                                                                                           cm long (L) as shown in Fig. 2 and spanning                                                                                                                    16.51 cm
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                                                                                                     .       WINDTUNNEL NOZZLEEXIT
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                                             0                                 -40                                       -30                                       -20                                 -I0                                  0                              I0                                                           20                                    30                         ,.0
                                                                                                                                                                                                                  x t
                                                                                                                                                Fig.               2 Slot              bleed experiment                                           schematic.

                     NAS A/TM _199                                    8-20697 4                                                                                                                                   2
                                                                                                          Table 1 W'md-tunnel                   operating      conditions.

 FROM   40 PSI                                            I                                 I

                                                                       SHOCK    GENERATOR

                                                                                                                                                        1.98                 2.46

                                                                                                                  Pt,0      kPa                     138.0                    172.4

                 ASME   CALIBRATED
                                                                                                                   Tt,0 °K                          293.0                293.0

                                                                                                              Re x 10 -7/m                              1.77                 1.75

                                                                                                                  t_re f    cm                          2.62                 3.06

                                                                  TO   VACUUM
                                                                                                                                                    0.565                0.712

                                 MASS-FLOW   PLUG                                                                                                   0.201                0.196

                                                                                                                     nine                           1.262                    1.260
Fig. 4 Slot bleed experiment                        wind-tunnel    schematic.
                                                                                                              Cs,_y         x 10 3                      1.50                 1.29
(S) as shown in Fig. 3. As indicated in Fig. 2, an oblique
                                                                                                            "'              kg/s/cm                0.04047              0.03768
shock-wave       is generated   by a rotatable and translatable                                            mbl_re]

(x-direction)     sharp-edged    plate (shock generator).    The
shock generator        spans the full width (30.5 cm) of the
                                                                                                    experience     14 has shown            that for both of these            cases the
wind-tunnel.      For a given deflection angle (c0, the shock
generator is translated until an inviscid shock-wave        orig-                                   shock strength           is sufficient to separate the wind-tunnel
                                                                                                    boundary-layer.           Hereafter, the two cases will be referred
inating at the generator leading edge will impinge on the
upstream      edge of the slot. For the purpose of setting                                          to by the M198A6              and M246A8            designations.
the axial position of the shock generator, the shock angle
(fl) is calculated     based on the core Mach number mea-
sured in the upstream reference plane indicated in Fig. 2.                                                                            Instrumentation
Boundary-layer    fences were used to isolate the wind tun-
nel test section comer flow from the interaction region.                                                  The experiment was instrumented       to measure surface
     The experiments were performed in the NASA Lewis                                               and flowfield smile pressure, flowfield Pitot pressure, and
Research Center 1X1 ft. Supersonic Wind Tunnel (SWT)                                                the total mass-flow through the slot. All flowfield mea-
which is a continuous-flow    facility with Mach num-                                               surements were made in the z=0 plane of symmetry using
ber variation provided by replaceable       nozzle blocks.     A                                    conventional     (intrusive) pressure probes. Because it was
schematic    of the 1X1 SWT experiment            is shown in                                       recognized that there would be large gradients and flow
Fig. 4. The approach boundary-layer       is the naturally oc-                                      angles in the slot, efforts were made to choose instru-
curring boundary-layer       on the wind tunnel wall.        The                                    mentation    least sensitive to these conditions.     However,
oblique shock-wave     interacts with the boundary-layer     and                                    as with all forms of intrusive type probes, inherent errors
a portion of the resulting distorted    boundary-layer     is re-                                   are introduced into the measurements    which are very dif-
moved through the bleed slot. The bleed flow exhausts                                               ficult to quantify. Regions of the flow that the authors
into a large plenum where the plenum static pressure is                                             believe to have a high level of uncertainty     are accord-
recorded.    The bleed flow rate was measured          with cal-                                    ingly identified.
ibrated ASME nozzles and the rate of mass-flow               was
                                                                                                         All pressures were measured       with a Pressure Sys-
controlled  by a choked mass-flow                       plug. Bleed vacuum
                                                                                                    tems Incorporated    electronic scanned pressure system us-
was supplied by lab-wide altitude                      exhaust and a 450 psi                        ing various range transducers.    All of the PSI transducers
air ejector system.                                                                                 have a manufacturers     quoted accuracy of _+0.07% of full-
       Data were obtained for two wind-tunnel       operating                                       scale span.
conditions.   These conditions are referred to by the wind-
tunnel core Mach number measured in the upstream ref-
erence plane (see Fig. 2). Table 1 summarizes       the wind-                                       Surface      Static Pressure
tunnel plenum condition       and the boundary-layer     char-
acteristics  measured  in the reference    plane for each of                                              Surface static pressure   data were measured      with
the reference   Mach numbers.     The unit Reynolds num-                                            pressure sensitive paint and with conventional  static pres-
ber reported in this table is based on the plenum condi-                                            sure taps. The pressure sensitive paint data were acquired
                                                                                                    using the technique described    by Bencic. 15
tions and the reference plane core Mach number (Mrel).
For the Mrey=l.98       condition, the shock generator de-                                               The tap data were measured with +_5 psi range pres-
flection angle was set at c_=6° and for the Mr_y=2.46                                               sure transducers  which yield an absolute   accuracy   of
condition, the deflection   angle was set at a=8 °. Previous                                        6P_--__ ,048 kPa (0.007 psi).

NASA/TMD1998-206974                                                                             3
                 C SLOT                                                                                     y
0.794                                              --__-0.254         THICK
                                1.588 0.254DIA
                            STAP(5PL)                           BRASSSHIM
                        ,   .      .


                  :::                   50.8

                  ,o,                                                                                                                       I        I

                   ,.:o.                                              1,0 DIAS.S.                                                           I        I
         ,=        =,,                                                                                                                      I        I
         ,,        ,,,
                  o:, o                                  I"           TUBING                                                                I        I:

         o,        ,,o                                                                                                                      I:       !
         ,,        o,,
                  o:,,                                                (5 PL)                                                                !
        o.        ,,.
                                                                                                                                            :        I
                                                                                                                                            °       _

        UU U                                                                            Fig. 6 Knife-edged          static pressure   probe        installation.

                                                                                        an inner-to-outer    diameter     ratio   of 0.4.        Based    on data
                                                                                        presented by Bryer and Pankhurst, 16 the critical angle for
   Fig. 5 Knife--edged                 static pressure        probe   :detail.          this probe configuration is +15 ° for incompressible    flow
                                                                                        and slightly larger for supersonic   flow.  Calibration    of
Flowfield         Static Pressure                                                       the probe in a Mach 2.0 stream showed that the probe
                                                                                        would read within 2% of the actual Pitot pressure    for
       Flowfield static pressure data were measured using
                                                                                        pitch angles as high as +_20° . The probe was actuated
a knife-edged     probe as shown in Fig. 5 and 6. For the
                                                                                        from the bleed plenum in the x and y-directions.
purpose of rneasuring      fiowfield static pressure, a _special
bleed plate insert was fabricated which had small grooves                                     The flowfield Pitot pressure   data were measured
machined _ong the vertic_ interior surfaces of the slot                                 with +15 psi range pressure transducers   which yield an
on the centerline      as shown in Fig. 6. These grooves                                absolute accuracy of 6Pt2---_- .145 kPa (0.021 psi).
serve as a guide for the knife--edged       probe whose details
are shown in Fig. 5.         Five 0.254 mm diameter          static                     Total Mass-Flow
pressure    taps are located     on a 0.254 mm thick sharp-
edged brass shim,         The taps are offset from the slot                                  The total mass-flow through the slot was measured
centerline    such that a reversal (180 ° rotation      about the                       with an ASME flow nozzle. The flow rate is determined
stem) of the probe in the guides will yield data between                                from a semi-empirical    relation which gives the mass-flow
the original tap positions     for a total of ten data points in                        as a function of nozzle geometry, temperature,       upstream
the x-direction.     The probe was zeroed with the center                               pressure    and the pressure     drop across the nozzle.      A
of the taps lying in the y=0 plane (wind-tunnel          surface).                      description    and uncertainty     analysis of this system is
A remotely controlled actuator in the bleed plenum then                                 given by Willis et al.n7 For the present measurements,
pulled the probe through the slot.          Knife-edged    probes                       the uncertainty   is estimated    to be between          +_.2.20% of the
                                                                                        calculated   mass-flow.
are known to be very sensitive toyaw            misalignment.      16
However, since all of the data taken were confined to the
plane of symmetry      and the machined                   guides prevented
aerodynamic   deflection   of the probe,                 the errors due to                                  Results    and Discussion
yaw misalignment      were minimized.
      Like the surface static pressure, the flowfield static                            Sonic Flow    Coefficient
pressure   data were measured     with +__5psi range pres-
sure transducers    which yield an absolute    accuracy    of                                 Numerical  simulations      of a bleed region utilizing a
6P---__ .048 kPa (0.007 psi).                                                           global bleed model must specify the bleed mass flow,
                                                                                        which usually varies with local flow conditions,            as a
                                                                                        boundary condition.    The mass flow may be determined
Pitot Pressure
                                                                                        from semi-empirical     relations     or from an experimental
   Pitot pressure measurements   were made with a 0.508                                 table look-up for a given bleed configuration.       The abil-
mm diameter square-edged   round Pitot tube probe having                                ity of a bleed configuration        to remove   mass from the

NASA/TM--1998-206974                                                                4
                                                                                                   the M246A8      case, respectively.     In these figures, the
       17.5                                                                                        lower half of the plot corresponds       to the no-bleed  case
                                                                                                   and the upper half corresponds      to the highest bleed flow
                                                                                                   rate attainable with the bleed system. The bleed rates are
       12.5                                                                                        given as a normalized    mass-flow defined as"

                                                                                                                                       100.         rh
X                                                                                                                       /Tt_                                                   (2)
o'                                                                                                                                     •     r}l/

         5.0                                                                                       where rh_,z,,e f is the unit boundary-layer   mass flow in
                                                                                                   the upstream reference plane (see Table 1)and S is the
         2.5                                                                                       spanwise dimension of the slot (see Fig. 3). While the                               _!!ili
         0.0                                                                                       data for each flow rate are shown for only half of the
               0        5       10      15        20          25       30        .55      40       bleed plate, results were obtained across the entire plate .....
                                      .        -       )   X 100                                   and exhibited a high degree of symmetry     about the wind
                                                                                                   tunnel centerline   (z=0).
       Fig. 7        Sonic    flow coefficient             distributions,        a=O.
                                                                                                        For the cases with no bleed flow (the lower half
                                                                                                   of Figs. 8 and 9), the results indicate a reasonably two-
    boundary-layer           is typically     quantified           by the sonic        flow
    coefficient:                                                                                   dimensional   flowfield except in the vicinity of the slot
                                                 rh                                                ends. As the bleed flow-rate is increased,         the upstream
                                Q=                                                       (1)
                                       l_ideal,chok,        ed                                     influence of the shock-wave    decreases, but not uniformly
                                                                                                   across the span of the slot. At the center of the slot
    which      is usually      presented      as a function            of the ratio       of
                                                                                                   (z=0), the upstream influence      is significantly    less than
    bleed plenum static pressure to ffeestream  total pressure•
                                                                                                   at the slot ends.     This spanwise    variation    of ups_eam
    Sonic flow coefficient distributions were measured for the
                                                                                                   influence indicates that more mass flow is removed           near
    case of an undistorted   approach boundary-layer   (a=0 °)
                                                                                                   the center of the slot. Although the interaction is not two-
    for four reference Mach numbers and are shown in Fig. 7.
                                                                                                   dimensional,   the results were observed    to be symmetric
    These data are presented    here for reference and can be
                                                                                                   about the wind tunnel centerline (z=0 plane) and are still
    used to validate flow coefficient models for use in CFD
                                                                                                   cogent for a three-dimensional   validation    case.
    methods.        A comprehensive          discussion            of flow coefficient
                                                                                                        Conventional      surface     static taps were located            axially
    behavior for the present slot and other                          bleed configura-
    tions is given by Willis et al. 17                                                             on the wind-tunnel    centerline    through the interaction     re-
                                                                                                   gion. Normalized     centerline    static pressure distributions
          When an oblique shock-wave      is positioned upstream
                                                                                                   for three bleed rates are shown in Figs. 10 and I 1 for •the
    of the slot, the local Mach number in the vicinity of the
                                                                                                   M198A6 and M246A8           cases, respectively.   The normal-
    slot is reduced   from the upstream      reference  value and
                                                                                                   ized wall pressure is defined as:
    a shift to a different flow coefficient     curve occurs.  An
    interesting feature in Fig. 7 is the presence of a distinct
    kink in all of the flow coefficient distributions. This is                                                           -      _ (Pw - P_,_)                                  (3)
                                                                                                                         Pw     -   (P3,i-          Pl,i)
    the location where choked flow occurs in the slot. The
    effective throat area, however, is determined    by the size                                   where i denotes     the inviscid        no-bleed         condition.   In these
    of the separation  region (see Fig. 1), which can vary with                                    plots, the symbols represent data from the conventional
    local flow conditions.     Lowering    the plenum pressure                                     surface pressure tapsand the solid lines represent data de-
    below this choke point causes an increase in effective                                         duced from the pressure sensitive paint. The theoretical
    throat     area and more mass            is passed           through     the slot.             inviscid distribution for the no-bleed  case is also repre-
                                                                                                   sented in these plots. The layout sketch at the top of the
    Surface        Static Pressure                                                                 figures shows the inviscid no-bleed wave structure.     The
                                                                                                   three bleed rates shown represent the no-bleed, maximum
          Pressure  sensitive    paint was applied                          to the bleed
                                                                                                   bleed attainable   for the configuration, and an arbitrary
    plate insert shown in Fig. 3. Initially,       it                        was thought           bleed rate between the two extremes.
    that with the boundary-layer       fences and                          the relatively
                                                                                                        From Figs.     10 and       11, the following               observations
    large span-to-width     ratio (S/D = 16.5) of                           the slot, that
                                                                                                   can be made:
    the interaction near the center of the slot would be nearly
    two-dimensional.      For the case of no bleed flow, this                                      1.   The upstream           influence        of the shock-wave   and
    was indeed the case. However, when the bleed flow was                                               boundary-layer         interaction       is reduced by approxi-
    non-zero, the interaction became more three-dimensional.                                            mately 1.5 cm for the M198A6 case and by approx-
    This behavior is illustrated  by the surface static pressure                                        imately 2.5 cm for the M246A8 case when bleed is
    obtained with the pressure sensitive paint and shown in                                             applied.  This is presumably  due to the removal                           or
    Figs. 8 and 9 which correspond     to the M198A6 case and                                           reduction of the separated region.

    NASA/TM--1998-206974                                                                       5
                                                                          " :   / "/   • :_      ,     •   • ....            ,:   • :    i           •:          ::• il i_ .i     : >::•L::#: ii/:_/             ;:: i: i    _          :: •:}% !!/• !!- !:   ! i ii :% !ig•ilL:•9:!!!:_ii:izli
                                                                                                                                                                                                                                                                   ¸                      _ _iil}'iii:! ii:iii_: iiii:i                                                                                    !


                                                              1.96                                                                                                                                                                           ' ' I .....                      I .....            .'1 '. '._'_                 I'

                                                                                                                                                                                                                                        "'.                                           ////
                                                                                                                                                                                                                                                 •                                /

                                                              1.85                                                                                                                                                                                   "...           ,,."

                                                                                            E        10                                                                                                REFLECTED/"                                                       "     .T.E. EXPANSION
                                                              1.73                          o
                                                                                                             "",,             LE.            SHOCK                   @                                 SHOCK,,                    "_'_                                            "'".

                                                                                                             -                    ,,           :......:      ..................
                                                                                                                                             .... ..................                                          :',/          __                                                                   "',,             ,
                                                                                                                         ',_?i: i;_::;:':;_:5_::;:;::i;:;:::i
                                                                                                                                                ;':;;:.;:;;;i:i_:g!:i:i!!iii!ii::ii!i:i:;i !i:.!;:!::;:•
                                                                                                                                                        ;;':i:                "i '_:;           :                                                                                            FENCE                : ",,




                                                                                                     1.0                                                                r                                                                                                     ......
                                                                                                                                                                        i¸                                                                                                                                            .....

                                                                                                                                                                        i               .........                      INVISCID                       NO-BLEED                          THEORY
                                                               1.03                                                                                                                     .        o                    r_=o
                                                                                                                                                                                                                       &= 5.o4%
                                                                                                                                                                                                                       _ = 7.63%

        Fig. 8    Surface     static pressure from                                                                                                                                      SYMBOLS                       - STATIC                       TAP DATA

       pressure   sensitive    paint, case M198A6.                                                                                                                              SOLID LINES-                                PRESSURE                          SENSITIVE                          PAINT DATA

                                                                                                                    ,I,,,,1,,,,                                                 ,,,,I                  ....                 1,,,,I,,_,1                                               ....              1 ....                I,

                                                                                                                  -10                          -5                      0                       5                      10                         15                          20                   25                         30
                                                                                                                                                                                                              × (o_)

                                                                                                                 Fig.                    10 Normalized                                                    centerline                                        (z=0)                       static
                                                                                                                                  pressure                         distributions,                                                 case M198A6.


                                               Q_                                                     20                                                                                        !''''                       I'          ''      ' I ''''                       I''''                    l''''l'
                                                                                                                                                                                                                                                                                                                       /,.     t

                                               EL              1.60                                                                                                                                                                                                REFLECTED                             ..-"

                                                                                            E                                                                               T.E. EXPANSIOI_"                                                                       SHOCK                     ...-."

                                                               1.39                                   10                                                                                                                                "-.                        ...

                                                               1.18                         >"                   I_SHOCK                                              Q                                                          .-'"                       "-...                                        C%



        Fig. 9 Surface    static pressure from
       pressure sensitive paint, case M246A8.                                                        1.0                                                                                                                                                                                                   ........
2.   For both cases, there is very little difference between
     the distributions corresponding     to the two non-zero                                                                                                             i               ........o                      r_=O
                                                                                                                                                                                                                        INVISCID                       NO-BLEED                              THEORY
     bleed rates.
                                                                                                                                                                                                                        I_ : 4.44%
3.   For the non-zero bleed cases, the pressure overshoots
                                                                                                                                                                        i!               __                             r_ = 6.38%
     the inviscid pressure distribution.  This overshoot in
                                                                                                                                                                                          SYMBOLS                      - STATIC TAP DATA
     pressure is due to the barrier shock (see Fig. 1) and,
     as expected, is greater for the higher bleed rate case.                                                        _!¢=ii.--_                            ....           i] SOLID LINES-                                    PRESSURE                          SENSITIVE                          PAINT                DATA

4.   The pressure sensitive paint data agrees quite well                                                                ,i        ....               1 ....                 1 ....                 I,,,,I,,,,I                                              ....                  I .....               I ....                     I,

                                                                                                                    -        10                 -5                      0                       5                       10                           15                      20                       25                     30
     with the conventional   static tap data.
                                                                                                                                                                                                                x      (cm)
5.   The drop in pressure below the inviscid value is
     due to the upstream influence of the expansion wave
                                                                                                                    Fig. 11 Normalized      centerline (z=0) static
     system set up by the experimental     hardware (shock
                                                                                                                       pressure distributions,    case M246A8.
     generator and fences).

NASA/TM--1998-206974                                                               6
Flowfield   Measurements Slotin.the
     Fortheflowfield                       onlyone
                          measurements, bleed             rate
was  considered   foreach    ofthetwooperating     conditions.
Thebleed    rateconsidered     corresponds   tothemaximum
attainable  withthebleed      system  andis thesame     asthe
         bleed condition
highest rate                    reported  forthesurface  static
pressure Withreference            tothesonic   flowcoefficient
distributions   (Fig.7), for all localMachnumbers,         the
highest  flowrate  attainable  withthe   bleed sytem isalways
totheleftoftheaforementioned pointintheflow
coefficient  curves.
     Theknife-edged probe static        wasused    tomeasure                                                                                                                                                        >.

thestaticpressure the slot. A totalof 210points
                  at                  s
weremeasured thelocationshown Fig.12. Re-     in
suits fromtheflowfield      static pressure measurements   for
theM198A6                       c      a
                andM246A8ases reshown Figs. 3      in       1
and14,respectively.     For•presentation    purposes, thedata
have extrapolated          totheslotsurfaces     bysetting the
pressure thesurface                 t
                             equalo theflowfield      pressure                                                                                                                                                           -2.5
atthefirst pointawayfromthesurface. hedatain      T                                                                                                                                                                                     M,.'-_1.98                                _'-6.0    (dog.)

theselotsarenormalized                              total
                               bythewind-tunnel pres-                                                                                                                                                                                   P,.o= 137.9        (kPa)                  (Pp,../P,.o)X   100= 6.300

sure           Q
     (P_,0). ualitatively,                      a
                               thetwocasesreverysimi-                                                                                                                                                                    --3.0
                                                                                                                                                                                                                                        rh =7.63%
                                                                                                                                                                                                                                        ,,,,               I    ....          I     ....             I   ....
lar. Although    undoubtedly               by
                                 smeared thepresence        of                                                                                                                                                                   -0.5                    0.0                0.5                   1.0               1.5
theboundary-layer     developing    ontheknife-edged     static                                                                                                                                                                                                             x/D
probe,  thepresence                     s        of
                      oftheinterior egment thebarrier
shock clearly      seen  bythelarge     pressure          (
                                                  gadient in-                                                                                                                                                             Fig. 13 Normalized    static pressure
crease) theupperightregion theflowfield.It is
         in             r              of                                                                                                                                                                            distribution ((P/Pt,o) x 100), case MI98A6.
notclear, owever,     whether    theshock   isattached ornot.
Outside   thisregion,  thestatic             is
                                   pressure fairlyuniform                                                                                                                                                                  0.5
throughout    theslot.
            0.5         _         .....                                    I......                               .!'" ....                         I         ....
                                      %%                                                         s

                        !%,                                                                          ,_"           INVISCID                  SHOCK
                                                                    ,_*,                                         FLOW



                                                                                      • • • •                          • m • _ •                                      ++++++:++!i:;!i!!_;++:+++++++:++:.++:++i++
                                                                                                                                                         ...... ...........                       :



                                                                                 __                                    __                                                                                i

                                   .........................................     _    _         N        _   _         _       1_   _(   _

          -1.5                                                '+                 __                                    __m
                                        .                :::::::::::::::::::::   __                                    _1_                                                                                               -2.0
                                       ....................                                                                                                   _::.:_?_,_:.:..:::.......
                               .........................                                                                                                    ....................
                                                                                                                                                                    :.::.        .
                                                                                                                                                                ::_.. ::.:.:.:........

            2. 0
                                                                               •      •         1_       •   •         •       •    •    N                                                                               -2.5
                              ...........                     ............

                               :                                                                                                                                 ...............................
                                                                                                                                                          ......................            ...

                                                                                                                                                                  ...............                                                       P,.o= 172.4        (kPe)                  (P,,,.,/P,.o)X 100=4.000
                                                                                                                                                       :ii:;ii!iiiiiiiiiiiii:iiiii:iiiiiiiii:!i+i+_                                            =6.38_.
          -2.5                                                 _!!i_ __                                                __
                                                                                                                                                                                                                         -.5.0     ....                    I,          ,,   _ t,,          _,      I,           , , ,
                                                                                                                                                                                                                             -.0.5                       0,0                0.5                   1.0                   1.5
          -3.0                    ,         ,          ,        _          I     _        ,          ,       ,     i       ,        ,    _   ,     I          ....

                 -0.5                                               0.0                                          0.5                             1.0                                           1.5                        Fig. 14 Normalized   static pressure
                                                                                                                                                                                                                     distribution ((P]et,0) x 100), case M246A8.

     Fig.         12              Static                               pressure                                   measurement                                                 grid.

NAS A/TM--1998-                                               206974                                                                                                                                          7
           1.5           '      '    '    I     '       '     '         I   '       '            '        I        '     '        '       [     '            '        '    I     '        '     '
                        .....             INVlSCID                THEORY                     (NO-BLEED)

           1.0                                                          I                                                                                                                                                                                                             ,_                _"             _ FLOW
                                                                                                                                                                                                                                        0.0 "                                    _:*+++++++++++++++++-...................................
                                                                                                                                                                                                                                                     ..............................................................                                                              _•_ .................
                                                                                                                                                                                                                                                                                                                                                          _v<;:< .................. _
                                                                        I                                                                                                                                                                            _::::_:__:__*<:__:_:::_:_:_{_:_:::,_:
                                                                                                                                                                                                                                                     :.                                          +++++++++++++++++++.                                                      .::::.:.      :.::.::..::,
                                                                                                                                                                                                                                                                                                                                                           ,:::.:.............. .............
                                                                                                                                                                                                                                                     .................................                                                                          ;i::i;ii:iii:il;y% >: i: :iil;•il
                                                                                                                                                                                                                                                                                                                                                         !!!#2:111_iii           i 15
                                                                    I                                                        Z_               STATIC                      TAP        DATA
           0.5                                                          |                                                                                                                                                                            ............ .......         .++++++++++++++++++. :::::;::_:,::::iii_!!iii!!i_:i_ii'/ii:.i!}!iiii_i!i!!!iiiii:@
                                                                                                                                                                                                                                                         :: ...............                              "• ••.::= ##///;U//##::IU
                                                                                                                                              PSP DATA
                                                                                                                                                                                                                                                                                  +++++++++++++++++++, %i_,ii:,iii_i_,i_,i:i:}:iii;i_:!:i_:_:iii:iii:_i_ili_:i:i
                                                                                                                                                                                                                                                     :::::::::::::::::::::::::::::::::::::::::::::::::::                           !!
                                                                                #                                            <>               KNIFE-EDGED
                                                                                                                                              PROBE DATA
                                                                                                                                                                                                                                                                     .......... +++++++++++++++++++.                                                                 •...............
           0.0                                                                                                                                                                                                                                                                       _I_:_;•
                                                                                                                                                                                                                                                     .i::_;::i::i:/::i:•ii:;/:_:_]::¢:ii:iiii:/:]_iiiii::]:_::::_ii                                                               :_?#::
                                                                                                                                                                                                                                                                                                                                                         ii:///:]ii:=#_;#£::/:_•_]#::::/: ]
                                                                                                                                                                                                                                                                   . :, ..:. .........                ++++++++++++++-_-+++-
                                                                                                                                                                                                                                                     .......... {:...:_                                                                               ,. _!_/!_!_,_!i_!!_/_._._._._._._:_;:_

         -0.5                                             i__
                  _iiii!iiiiiiiiiiiiiii!iiiiiiiiiiiiiii!iiiiiiii!iiiii!iiiiiiii!iiiii_                                                                                                                                                                                                                                                                                   ::::::::::::::::::::::: •
                                                                                                                                                                                                                                                                                                                                                            :; _:_:_:::_#_

                                                                                                                                                                                                                                                                                                      ++++++++++++++-H-_++.                                _                          _i:i
                                                                                                                                                                                                                                                                                                                                                          :: ,i;:::i;:::::;:::::i;iii:;:i;}_i:i:i;i7
                                                                                                                                                                                                                          123                                                                                                                                          .............

                                                                                                                                                                                                                                                                                                      ++++++++++-+-+++-t-÷+++.                                                    :.::..-;
                                                                                                                                                                                                                                                                                                                                                          ::::_::_:_:_.:_:_:::_: ..... ,
                                                                                                                                                                                                         4                >...

                                                                                                                                                                                                                                   - 1.5                  >:_:_:_:_:_>:_:_:_
                                                                                                                                                                                                                                                      ;:_:_:;ii_i_ii +++++++++++++++++++.
                                         -2                        0                                  2                                   4                                6                             8                                                             .................

                                                                                        x (0m)                                                                                                                                                   !   ........                                         ++++÷++++++-a-++++++.                                                                  :...........
                                                                                                                                                                                                                                                                                                                                                               ...................................... i

                                    a) Case                   M198A6,                                     rh=7.63%.
                                                                                                                                                                                                                                   - 2.oi                                                             +÷÷÷÷÷÷._÷÷_÷÷÷÷÷
           1.5           '      '    '    I     ....                I       '       '        '        I            '     ,        ,       I     '        '            '    i     '    '         '                                                                                                     ÷+++++++-H-+++++++++.

                                          INVlSClD                THEORY                     (NO-BLEED)                                                                                                                                              ...................       _:_:_.:_;::.:..::::ii::+++++++_-+.+_++++++++.

                                                                                                                                                                                                                                                                     ......i!...... :: :.-I-+++++++-H--i-++++++++.                                    .........................              i        .

           1.0                                                                                                 -                      _              .           ......                              :

                                                                        I                                                    A                STATIC                      TAP        DATA
           0.5                                                                                                                                                                                                                     -3.0
                                                                                                                                              PSP                DATA
                                                                                                                                              KNIFE-EDGED                                                                              -0.5                                                0.0                             0.5                     1.0                                           1.5
                                                                            o                                                <>               PROBE                       DATA

                                                                                                                                                                                                                                 Fig.      16                Pitot                          pressure                         measurement                                grid.

         -0.5                                                                                                                                                                                                    separation                in the slot which                                                          severely            reduces                 the effective
                                                                                                                                                                                                                 flow      area.           The                    data                      do                 indicate,           however,                   that                     the                  flow

         -1.0                                                                                                                                                                                                    reattaches                before                          exiting                              into     the        bleed          plenum.

                 -4                      -2                        0                                  2                                   4                                6                                              The           Pitot             and static                                           data     were           combined                        to calculate
                                                                                        x (°_)                                                                                                                   the     local            Mach                       number                                     distribution                in the                 slot.                              If the

                                    b) Case                   M246A8,                                     _=6.38%.                                                                                               Pitot-to-static                      pressure                                        ratio            (Pt2/P)             is less            than                      or equal
                                                                                                                                                                                                                 to 1.893               (subsonic                                 flow),                        then         the Pitot             pressure                                is equal
                      Fig.           15         Normalized                                           centerline                                static                                                            to the           total       pressure                                      and                 the     Mach              number                is calculated
                 pressure                     distributions                                          in the                       y=0               plane.                                                       from        the         isentropic                                       relation:

           The        normalized                            static          pressure                                   distributions                                       from               the
the       surface
                                                in the y=0
                                                pressure                        upstream
                                                                                        plane                          are plotted
                                                                                                                             and              downstream
                                                                                                                                                                    along            with
                                                                                                                                                                                                                                                                                               (       1+
                                                                                                                                                                                                                                                                                                                   7 - t M2

the       slot     in Fig.                    15.           The             upstream                                    influence                                  of the                     slot
                                                                                                                                                                                                                 If the Pitot-to-static                                               pressure                         ratio       (Pt;/P)               is greater                                         than
expansion                    can         be         clearly                 seen                     in                the            pressure                             sensitive
                                                                                                                                                                                                                 1.893           (supersonic                                   flow),                           the Mach                  number                is calculated
paint       data.            The          low               pressure                     in the                          upstream                                  half          of the
                                                                                                                                                                                                                 from.the               Rayleigh-Pitot                                                         tube      equation:
slot       opening                  indicates                     a stalled                               condition                                 and                   that       very
little      flow             is being                  passed•                  through                                this            region.

           A 0.508                  mm          diameter                        Pitot                         probe                   was            used                      to mea-
sure       the Pitot                 pressure                     in the                  slot.                        A total                      of 441                       points
were        measured                      at the              locations                                   shown                           in Fig.                         16.        Nor-
realized           Pitot            pressure                      distributions                                         for the                     M198A6                                    and
                                                                                                                                                                                                                  For     purposes                       of the calculation,                                                      the local              static                         pressure
M246A8                  cases             are           shown                    in Figs.                                    17           and                    18,           respec-
                                                                                                                                                                                                                 (P) data               were             interpolated                                             onto           the      data      grid                used                          to ac-
tively.           Near              the        top           of the                     slot,                  very                   large                      flow            angles
                                                                                                                                                                                                                 quire           the Pitot                 data.                           Recall                     that       the      static      pressure                                            at the
relative          to the Pitot                         probe                stem                     are expected                                    and the                         Pitot
                                                                                                                                                                                                                 slot    surface-was                                 set equal                                   to the           first     column                     of data and
data       should               be considered                                   very                  uncertain.                                    This                  probably
                                                                                                                                                                                                                 recognize                that           this may                                       introduce                 errors         in the Mach                                              num-
accounts               for the                 lack           of a discrete                                              barrier                     shock                       in the
                                                                                                                                                                                                                 ber     values            nearest                              the slot                          surface.                The      calculated                                             Mach
data.           Also,           it should                     be noted                                that                   the          Pitot                     probe                 will
                                                                                                                                                                                                                 number             distributions                                           for the M198A6                                  and M246A8                                                    cases
sense           nearly              static             pressure                         in            regions                             of        reverse                          flow.
                                                                                                                                                                                                                 are shown                 in Figs.                               19 and                          20, respectively.                          The                     presence
These           results             indicate                 two fairly                               extensive                                regions                         of flow

N AS A/TM                       _1998-                 20697                4                                                                                                                                8
                                                                                                                                               0..51         .......                I''''                     [,"'''                      I .....



                                                                                                                                                             ,.,:_._                                            _=_.o(_._')
                                                                                                                                                             P,,o=137.9             (kPa)                       (P       /P,,o)X100=6.300

                                                                                                                                                             _n =7.63%

                                                                                                                                             -3.0     ,,,,                         I,,,,                 I,,,                      ,. I ,. , , ,
                                                                                                                                                 -0.5                            0.0                   0.5                           1.0         1:.       5


                                                                                                                                                Fig.           19           Calculated                        Mach              number
                                                                                                                                                        distribution,                          case             M198A6.

                                                                                                                                               0.51          .....                 I'        ' ' ' I,"                  ' '               I''       ' '
                                                                                                                                                      _*%,                                      /        : INVISCID               SHOCK

                                                                      ;::+: .
                                                            ................:: .....................................

                                                            i i i:iii_iiiiiii_ilili_i:iiiiiiiiiii:iii_i_iiiiiii!iii_iiiiiiiiiiii:
                                                                                                         •                                                                                    :#

                                                            _ ::iii%ili:i:i:i_ili_!!i:i:i;:i_i       "

      £3                                                                                                                                E3


                                                            i !}iiii!iiiiiiiiii:iiiiiii:iii               !iiii_

                                                            ili iii:i:i:i_i:i:iiiiiii_i!i:i:i:i:i:!:i:i_i:iii:i:i:i
                                                            i il i iii:i:i:ii_ili:i
                                                                              iiiii:iiiii!i:ili:ii!:ii         ii
                                                                                                                                             -.o                                   NNI i)

                                                              i                                     iii:"

                                                                                                                                             - 2.5                                  _iii!iiiiiiiiiiiii!iiiiiiii

                                                                                                                                                             M,.,=2.46                                          ==8.0          (deg.)

                                                                                                                                                             P,,o=       172.4      (.kPa)     .....   ....     (P:#../P,,o)XIO0=4.000

                                                                                                                                                      & =:6.38_
                                                                                                                                             -3.0     ,,   ,,   I,,,                                   _ I,,,,                            I .....
                                                                                                                                                 -0.5         0.0                                       0.5                             1.0               1.5

             Fig.    18    Normalized        Pitot    pressure                                                                                  Fig.           20           Calculated                        Mach              number
      distribution        ((et2/et,0)   x   100),    case     M246A8.                                                                                   distribution,                          case             M246A8,

NAS    A/TM--1998-206974                                                                                                            9
of the separation  region on the upper left surface of the                                                                From the surface static pressure distributions     (Figs. 8
slot causes an aerodynamic   convergent-divergent     nozzle                                                      and     9), we inferred that the mass-flow       distribution     in
effect.   As the flow expands downstream        of the aero-                                                      the spanwise (z) direction was not uniform,   but passed
dynamic    throat, the Mach number increases       to a peak                                                      a higher mass-flow at the center of the slot. If we as-
of about M=2 for both cases. The back-pressure          in the                                                    sume         that the total temperature                                                              of the flow in the slot is
plenum, however, is not low enough to maintain super-                                                             the same as the wind-tunnel   plenum, then in conjunction
sonic flow. But rather than shocking     down to subsonic                                                         with the static pressure and Mach number distributions    in
flow as would happen for the mviscid case, the plenum                                                             the slot, and assuming the ideal gas law applies, we can
pressure feeds up through the boundary-layer      on the slot                                                     calculate a mass-flux (-pV)   distribution in the slot. In-
surface which causes the flow to separate. Compression                                                            tegrating the -pV distribution   along lines of constant (9)
waves off of this separation    act to decelerate    the flow                                                     locations and then dividing by the unit mass-flow       in the
gradually but still the flow exits the slot supersonically.                                                       reference  boundary-layer   will yield a normalized     mass-
The upper separated region also leaves a rather large sub-                                                        flow in the plane of symmetry:
sonic wake which accelerates     to a near sonic condition
at the slot exit.

      Following  Kline, Is the propagation   of the uncer-                                                                                                                                                      f -pVdx
tainty in the measured Pitot and static pressures into the                                                                                                  _c_-                              100.              o                                                                        (7)
calculated Mach number was estimated from the follow-                                                                                                                                                                      bt,r:e]

ing equation:
                                                                                                                  There        are,        however,                                      at least                    three               sources                 of errors                   to
                                                                                                                  consider when performing    the integration:  regions of high
                                                                                                                  flow angle relative to the Pitot probe stem, regions of
         6M       -      --_6P                        +        OPt26Pt2                           (6)             reverse flow, and regions where the static pressure has a
                                                                                                                  high gradient near the wall which makes our extrapolation
                                                                                                                  assumption   uncertain.  The integrations    were performed
where 6P, 6Pt2, and 6M represent       the uncertainty in
                                                                                                                  along the 21 rows of data and the results are shown in
the static pressure, Pitot pressure, and Mach number,
                                                                                                                  Fig. 22. In this plot, the centerline mass-flow  calculated
respectively.  Using equation (6), the uncertainty in the
                                                                                                                  from equation (7) is normalized      by the bulk mass-flow
Mach number was evaluated at each point in the flowfield
                                                                                                                  measured with ASME nozzle and presented as a function
and the results are shown in Fig. 21 as a function of
                                                                                                                  of 9 location. From this figure we can estimate that for the
Mach number.     Note that this uncertainty   includes only
                                                                                                                  bulk mass-flows considered,    on the average the centerline
the pressure measurement   uncertainty   and not errors due
                                                                                                                  mass-flow is roughly 50% higher than the bulk mass-flow.
to probe    interference                      The results indicate
                          or interpolation .....
                                                                                                                  Also the increase in centerline mass-flow through the slot
that the   uncertainty   becomes excessive       for a local Mach
                                                                                                                  indicates a significant spanwise convergence     of the flow
number     less than 0.5 which is confined to relatively small
                                                                                                                  within the slot.
regions    of the flowfield.


                                                                                                                                                                         I        ....                      I          "    '        '    I   '       ....            I   ....
  0.25        ....                 I                       I                                                              .0          i    ....

              D_                                                                                                                                                                                                                                                                               )
  0.20                                                                                                                  1.8

                                               o      Case M198A6 ( rh = 7.63%)
  0.15                                                                                                                  1.6      -        OOOOO                                          O

                                               D      Case M246A8 ( rh = 6.38%)                                                  •"C] []           []       D       []       []          []        O
                                                                                                                 I'E                                                                                        o o o o

                                                                                                                                               o            Case                      M198A6                    (N-                 7.63%)
   0 O5                                                                                                                 1.2
                                                                                                                                               []           Case M246A8 (_ = 6.38%)
                                       - _-'.'-7.._            -   "_   ....   _   ._    _

   0.00                   ,   Ot                                                              ,   ,
                                                                                                                          .0          I    ,        i   l       .        I        ,      ,     ,       ,    I    ,     ,        ,    ,    I       ,   ,      ,    i   !   I      i   i   i

       0.0                    0.5                         1.0                           1.5           2.0                       -2.5                        -2.0                                   -1.5                             -1.0                         -0.5                    0.0
                                                           M                                                                                                                                                    y/g

                     Fig. 21 Uncer_inty in Mach                                                                                                Fig. 22 Integrated   mass-flow
                     number versus Mach number.                                                                                                  in the plane of symmetry.

NASA/TM--1998-206974                                                                                        10
                    Concluding                                             8 Hahn, T. O., Shih, T. I.-P., and Chyu, W. J., "Numerical
                                                                           Study of Shock-Wave/Boundary-Layer        Interactions with
    Developmentglobal        bleed         for
                                   models boundary-                        Bleed," AIAA Journal, Vol. 31, May 1993, pp. 869-876.
     c      will         e             and
layer ontrol requirexperimental computational
                                                           9Rimlinger,    M. J., Shih, T. I.-P., and Chyu, W. J.,
synergism                         of
          dueto thecomplexity theflowfield.The
flowfield      a
         inside bleed    slotused control noblique
                                   to         a            "Three-Dimensional      Shock-Wave/Boundary-Layer       Inter-
shock-wave andturbulent    boundary-layer interaction has  actions with Bleed Through a Circular Hole," AIAA Pa-
                                                           per 92-3084, July 1992.
been studied experimentally.                     we
                                Fromthis study can
drawthefollowing    conclusions:                           l°Chyu, W. J., Rimlinger, M. J., and Shih, T. I.-P., "Ef-
                                                           fects of Bleed-Hole     Geometry   and Plenum Pressure on
1. For the configuration       tested,despitea two-
                                                           Three-Dimensional       Shock-Wave/Boundary-Layer/Bleed
    dimensional  flowfield                   case,
                           forthezero-bleed appli-
    cation ofbleed-flow   resulted athree-dimensional Interactions," AIAA Paper 93-3259, July 1993.
              S         s
    flowfield. urfacetatic    pressure indicate
                                        data          that 11 Shill, T. I.-P., Rimlinger,   M. J., and Chyu, W. J.,
    theslotpasses mass-flownear          thecenter   than "Three-Dimensional       Shock-Wave/Boundary-Layer       Inter-
    attheends. hismay       verywellbethecase ac-  in      actions with Bleed," AIAA Journal, Vol. 31, No. 10, 1993,
    tualinletbleed   systems.                              pp. 1819-1826.
2. Theflowfieldin the slotis characterized the    by       12Rimlinger,    M. J., Shih, T. I.-P., and Chyu, W. J.,
    presence              s
             ofabarrier hock nd  a twolarge    separation 'q'hree-Dimensional      Shock-Wave/Boundary-Layer       Inter-
    regions.These              serve
                      featm'es to reduce        theflow    actions with Bleed Through Multiple Holes," AIAA Pa-
    coefficient thebleedpassage         andtheireffects per 94-0313, Jan. 1994.
    should              in
           beincluded flowcoefficient       models.
3. Dueto anaerodynamic effect f thesepara- 13Harloff, G. J. and Smith, G.AIAA "On Supersonic-Inlet
                             throat        o               Boundary-Layer     Bleed Flow,"         Paper 95-0038, Jan.
    tion,much oftheflowthrough      theslotissupersonic. 1995.
    Thedata             in             a
             presentedthisreport reavailable        from   14porro,    m. R. and Hingst,       W. R., "Use of Sur-
thefirst author n magnetic            or
                              media via theinternet face Heat Transfer Measurements as a Flow Separa-
(fsdavishopi.lerc.nasa.gov).                               tion Diagnostic in a Two-Dimensional      Reflected Oblique
                                                                           Shock/Turbulent  Boundary-Layer    Interaction,"   AIAA   Pa-
                                                                           per 93-0775, Jan. 1993.
                                                                           15Bencic,    T., "Experiences Using Pressure Sensitive
1Delery,                           Boundary-Layer                          Paint in NASA Lewis Research Center Propulsion     Test
Interaction ndIts Control,"Progress in Aerospace                           Facilities," AIAA Paper 95-2831, July 1995.
Sciences,   Vol. 22, No. 4, 1985, pp. 209-280.
                                                                           16Bryer, D. W. and Pankhurst,      R. C., Pressure-Probe
2Hamed,     A. and Shang, J. S., "Survey      of Validation                Methods for Determining   Wind Speed and Flow Direction,
Data Base for Shock-Wave/Boundary-Layer         Interactions               Her Majesty's  Stationary Office, London, England, 1971.
in Supersonic   Inlets," Journal of Propulsion and Power,
                                                                           17Willis, B. P., Davis, D. O., and Hingst, W. R., "Flow
Vol. 7, No. 4, 1991, pp. 617-625.
                                                                           Coefficient  Behavior   for Boundary-Layer  Bleed Holes
3 Hamed, A. and Lehnig, T., "An Investigation  of Oblique
                                                                           and Slots," AIAA Paper 95-0031, Jan. 1995.
Shock/Boundary-Layer/Bleed    Interaction,"   Journal    of
                                                                           18Kline, S. J., '"rhe Purposes of Uncertainity         "
                                                                                                                          An alysls, "
Propulsion and Power, Vol. 8, No. 2, 1992, pp. 418-424.
                                                                           ASME Journal of Fluid Engineering,   Vol. 107, June 1985,
4Hamed,     A., Shih,    S., and Yeuan,      J. J., "A Parametric
                                                                           pp. 153-160.
Study of Bleed in Shock/Boundary-Layer              Interactions,"
AIAA Paper 93-0294, Jan. 1993.
5 Hamed, A., Yeuan, J. J., and Shih, S., "An Investigation
of Shock Wave/Turbulent       Boundary-Layer    Interaction
with Bleed Through Normal and Slanted Holes," AIAA
Paper 93-2155, June 1993.
6Hamed,     A., Yeuan,    J. J., and Shih,     S., "Flow    Charac-
teristics in Boundary-Layer       Bleed      Slots with    Plenum,"
AIAA Paper 95-0033, Jan.         1995.

7 Hamed, A., Yeuan, J. J., and Shih, S., "An Investigation
of Shock Wave/Turbulent       Boundary-Layer    Interaction
with Bleed Through Slanted Slots," AIAA Paper 93-2992,
July 1993.

NASA/TM--1998-206974                                                  11
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 1.     AGENCY          USE   ONLY   (Leave    blank)              2.   REPORT      DATE                        3.   REPORT     TYPE   AND          DATES.     COVERED

                                                                                 April 1998                                             Technical            Memorandum
 4.     TITLE     AND    SUBTITLE                                                                                                              5.     FUNDING         NUMBERS

           Flowfield Measurements      in a Slot-Bled                        Oblique       Shock-Wave       and Turbulent
           Boundary-Layer   Interaction
 6.     AUTHOR(S)

           D.O. Davis, B.E Willis, and W.R. Hingst

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

          National Aeronautics and Space Administration
          Lewis Research Center
          Cleveland,          Ohio    44135 - 3191

 9.     SPONSORING/MONITORING                   AGENCY         NAME(S)        AND   ADDRESS(ES)                                                10.     SPONSORING/lVlONITORING
                                                                                                                                                       AGENCY          REPORT        NUMBER

          National Aeronautics and Space Administration
          Washington,  DC 20546-0001                                                                                                                   NASA TM--1998-206974

 11.     SUPPLEMENTARY               NOTES

          Prepared for the 33rd Aerospace Sciences Meeting & Exhibit sponsored by the American Institute of Aeronautics                                                                                 and
          Astronautics, Reno, Nevada, January 9-12, 1995. Responsible  person, D.O. Davis, organization code 5850,
          (216) 433-8116.

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

          Unclassified           - Unlimited
          Subject        Category:      02                                                      Distribution:        Nonstandard

          This publication is available from the NASA Center for AeroSpace Information, (301) 621-0390.
 13. ABSTRACT(Maximum200 words)
          An experimental   investigation was conducted to determine the flowfield inside a bleed slot used to control an oblique
          shock-wave   and turbulent boundary-layer   interaction. The slot was oriented normal to the primary flow direction and had a
          width of 1.0 cm (primary flow direction), a length of 2.54 cm, and spanned 16.5 cm. The approach boundary layer
          upstream of the interaction was nominally 3.0 cm thick. Two operating conditions were studied: M=1.98 with a shock
          generator deflection angle of 6 ° and M=2.46 with a shock generator deflection angle of 8 °. Measurements         include surface
          and flowfield static pressure, Pitot pressure, and total mass-flow through the slot. The results show that despite an initially
          two-dimensional    interaction for the zero bleed-flow case, the slot does not remove mass uniformly in the spanwise direc-
          tion. Inside the slot, the flow is characterized  by two separation regions which significantly   reduce the effective flow area.
          The upper separation region acts as an aerodynamic      throat resulting in supersonic flow through much of the slot.

 14.     SUBJECT         TERMS                                                                                                                                  15.    NUMBER        OF      PAGES

          Supersonic          flow; Boundary             layer control                                                                                          16.    PRICE    CODE

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         OF     REPORT                                        OF    THIS    PAGE                                OF   ABSTRACT

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