Effects of under expanded jet on afterbody drag of an axisymmetric configuration

Mathur, NB (1984) Effects of under expanded jet on afterbody drag of an axisymmetric configuration. Technical Report. National Aeronautical Laboratory, Bangalore, India.

	PDF tr_fm_tm_84_2.pdf Restricted to Repository staff only Download (3MB)
	Indexer Terms (Generate index codes conversion from application/pdf to indexcodes) indexcodes.txt Restricted to Repository staff only Download (4kB)

Abstract

Experimental investigations were carried out in 0.3m trisonic wind tunnel at NAL to study the effects of an underexpanded jet on afterbody/nozzle drag of an axiswrnetric configuration. Detailed pressure measurements were obtained on the boat tailed aferbody, of this model which accornmo6ated a convergent nozzle.Wind tunnel tests were conducted i n the free stream Mach number range of 0.5 to 1.1, keeping model at zero degree of incidence-13; Jet pressure ratio (Pm/Pd,) was also varied from 1.0 (Jet-off) to about 5.5 during this investigation. The jet at nozzle exit was sonic throughout thzse tests. Test results show that for free stream Mach nurnbers lower than jet Mach number ( ie.1.0) afterbody, boattail and base13; pressure drags increasad 2s the jet pressure ratio is increased from 1.0 (Jet off) to abut 3.0 but these drags decreased when the jet pressure ration is increased beyond 3.G.For higher free stream mach numbers (gt; 1.0) the boattail static pressures increased and afterbody, boattail an,l base pressure drags decreased as the jet13; pressure ratio i s increased beyond 1.0 (jet off). Results also show increase in afterbody pressure drag as the free stream Mach nunber is increased to about 1.0 and it decreased as the free stream Mach number is increased beyond 1.0. Schlcren pictures clearly indicate the movenent of the shock downstream of the boattail Shoulder as the free stream Mach number is increased above 0.95 and also the movement of shock upstream of the nozzle exit plane when the jet pressure ration is increased from 1.0 ( jet off) to about 5.5. Oil flow patterns also indicate the existence of the separated flows regions 0n the afterbody that occurs due to the shock-boundary layer interactions in the transonic Nach number range of this investigation. Location of shock on the afterbody as obtained from trailsonic calculations compares quite well with that obtained from the present experiments.nature of static13; pressure distribution compares well with the available NASA test results.

Item Type:	Monograph (Technical Report)
Uncontrolled Keywords:	Trisonic wind tunnel;Aircraft configuration
Subjects:	AERONAUTICS > Aerodynamics
Depositing User:	M/S ICAST NAL
Date Deposited:	14 Sep 2006
Last Modified:	24 May 2010 04:20
URI:	http://nal-ir.nal.res.in/id/eprint/2721

Actions (login required)

View Item