Environmental Engineering Reference
In-Depth Information
calculation of the
flight envelopes involves expertise from several disciplines
including gas turbine engine performance, aerodynamic, and structural analysis of
the vehicle of interest. The database in these areas are known accurately only by
OEMs followed by continuous re
fl
nement as the integration of the design proceeds
for particular application. However for academic exercise, one can use commer-
cially available software (viz. GasTurb
™
) and a vehicle with speci
ed
fl
flight
dynamic characteristics to create a
fl
flight envelope. The process of creating an actual
fl
flight envelope is beyond the scope of this article except to make the following
observations.
Perhaps the weakest link in the process of
finalizing the
fl
flight envelope is the
combustion designer
'
s limited analytical capability for
“
predicting
”
the ignition,
fl
flame propagation and acceleration, and lack of accurate information about the
engine conditions during the conceptual and preliminary design phases of opera-
tion. Therefore, the initial phase of estimates based on empirical database is fol-
lowed by extensive experimental evaluation and re
flight envelope.
Engine start conditions are very demanding including the available combustor
pressure drops (
nement of the
fl
Δ
P) an order of magnitude smaller than design
Δ
P and resulting
corrected
flow (Wc) less than one third of design Wc. This combined with lower
combustor inlet pressure and temperature as the airplane climbs poses enormous
challenge for achieving good spray atomization and combustion ef
fl
ciency and
thereby places limit on the maximum altitude ignition capability.
The stagnation pressure and temperature upstream of the engine inlet is straight
forward as given by one-dimensional gas dynamics relations, namely the normal-
ized ratios of stagnant temperature and pressure, respectively:
c=ðc
1
Þ
T
0
T
¼
c
1
2
M
p
0
p
¼
c
1
2
M
2
2
1
þ
1
þ
;
c heats at constant
pressure and volume. The process of determining windmill conditions even for a
gas generator is very complicated and only the engine core test in a full-scale test
facility provides a reliable data on P
3
and Wa
3
(Stearns et al.
1982
). The engine
(comprised of fan, low-pressure turbine, core, and accessories) and nacelle system
has to be installed in a
Here M is the
fl
flight Mach number and
ʳ
is the ratio of speci
flying test bed to provide quantitatively accurate conditions
for P
3
, T
3
, and Wa
3
. This
fl
fl
flight data is used to estimate fuel
fl
flow rate (Wf) for
ignition and
fl
flame propagation from empirical curves based on full-scale rig data.
s use proprietary empirical curves relating lean blowout and ignition fuel/
air ratios to combustor operating conditions. Experimental values of ignition and
lean blowout fuel
OEM
'
flow rates acceptable at the engine system level are determined
through an extensive tradeoff effort leading
fl
finally to establishing operating limits.
Here we show the one reproduced from Bruce et al. (
1977
), namely Fig.
21
which
will vary with different engines, their models, and vehicles. The maximum relight
altitude is also determined from the
flight test bed followed by some corrections
later because of the differences in the nacelle and airplane interaction. For this
fl
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