<DOC>
<DOCNO>
794
</DOCNO>
<TITLE>
experiments with a tapered swept-back wing of warren 12 planform at mach
 numbers between 0.6 and 1.6 .
</TITLE>
<AUTHOR>
hall, i.m. and rogers, e.w.e.
</AUTHOR>
<BIBLIO>
a.r.c. r + m 3271, part ii, july 1960 .
</BIBLIO>
<TEXT>
the development of the flow pattern on a wing of aspect ratio 2 828,
taper ratio 0 333, leading-edge sweepback 53 5 deg and 6 per cent
thickness/chord ratio in the streamwise direction has been described in
part 1, which discussed oil-flow patterns obtained on the surface of the
 wing . the complete programme of tests also included pressure plotting
at four spanwise stations and force measurements . these are discussed
in relation to the flow development in this part of the report .
the wing was tested at mach numbers between 0 6 and 1 6 for incidences
up to about 14 deg . the tunnel stagnation pressure was held constant at
a value near atmospheric pressure during the tests, so that the reynolds
 number varied with mach number .. at m 1 0 it was 2 3 x 10 based on the
 mean aerodynamic chord . boundary-layer transition was fixed by a
roughness band at the leading edge .
a detailed analysis has been made of the pressure distributions on the
surface of the wing and the chordwise distributions integrated to
determine the spanwise loading . the overall lift and pitching moment of
 the wing were also obtained from these data, as well as from direct
measurements using a strain-gauge balance, by means of which the wing
drag was also determined . these results are considered in some detail
to illustrate the effects of mach number and incidence on the flow about
 the model . a preliminary analysis is also made of the conditions for
boundary-layer separation due to shock waves on the wing surface . the
principal factor appears to be the component of mach number normal to
the shock front .
</TEXT>
</DOC>
