are commonly used in military aircraft. They have not used in civil aviation or for transport due to
several operational reasons. There are different concepts, which are being
proposed for supersonic civil transport. One of the concepts that presented in
this paper is Variable Cycle Engine (Adaptive Cycle Engine). Variable cycle
engines would be able to vary their bypass ratios, for optimum efficiency at
any combination of speed and altitude within the aircraft’s operating range, despite
traditional engines with fixed airflow.
supersonic highly loaded high pressure (HP) compressor was designed for a VCE
in the conditions of both single and double bypass modes in accordance with the
similarity principle. The blade profiles were designed by means of NACA
airfoil. Then, 3D numerical simulations were performed on the HP compressor of
both working conditions with different thermodynamic cycle parameters to
confirm the design methods and results. The one equation turbulence model of
Spalart-Allmaras was applied to solve Reynolds’s averaged Navier-Stokes equations.
The results of simulation indicate that the compressor performances are
satisfactory in both working conditions with high efficiency. Further research
reveals, wave structures in the supersonic compressor, behaviour of tip
clearance flow and the phenomenon of transition flow in boundary layer.
A Variable Cycle Engine (VCE) can be defined as one
that operates with two or more thermodynamic cycles. It is a type of aero
engine whose thermodynamic cycle can be adjusted by changing some components’
shape, size or position, and the cycle parameters, such as pressure ratio, mass
flow, bypass ratio and thrust. It can be varied between those of a turbojet and
a turbofan, making it to combine the advantages of both. These measures may
enable the engine to obtain the optimal thermodynamic cycle, and to acquire the
good adaptability to various flight envelopes.
The engine can work as the turbojet when the
aircraft requires high specific thrust, such as take-off, acceleration and
supersonic cruise. It also can work as the turbofan when the aircraft requires
low fuel consumption, such as
and subsonic cruise. The most important advantage expected from using VCE in
future supersonic transport is a substantial range improvements as compared to
a conventional engine. These range improvements are mainly achieved by reducing
the subsonic specific fuel consumption by around 15% (relative to a Turbojet)
and improving the fuel consumption at off-design by the extensive use of
variable geometry. The future VCE will have a low emission combustor and
afterburner. The noise level at take-off will be met by FAR part 36 requirement.
In other words, the future VCE will be environmentally accepted 2.
disadvantages are mainly an increase in the engine weight and a more complex
control system, therefore the reliability of the engine will be affected. The
performance of any VCE depends critically on the attainment of the predicted
technology level improvements. The purpose of research on VCE is to improve
off-design performances, in order to satisfy the needs of broad flight envelope,
large combat radius and long cruise duration 3.
work mode of VCE discussed in this article is presented in Figure 2:
Single bypass mode: The selector valve
is closed and all air goes through the Core Drive Fan Stage (CDFS). The fan
bypass flow bypasses the core engine through the inner bypass duct and remixes
with the core flow downstream of the low pressure turbine. The nozzle is full
open to shift the loading to the HP shaft to cope with the added work of the
CDFS. At the same time, the expansion ratio and the flow rate rise to increase
the specific thrust with low bypass ratio under supersonic and acceleration
Double bypass mode: The selector valve
is full open and the nozzle is now closed to unload the HP turbine and load the
LP turbine. The bypass ratio increases for best specific fuel consumption for
subsonic cruise and best exhaust velocity conditions for improved noise
suppression on take-off 4.
Axial flow compressor is one of the most important
parts of Gas turbine engine. Axial-flow compressors are used in medium to large
thrust gas turbine and jet engines. The compressor rotates at very high speeds,
adding energy to the airflow while at the same time compressing it into a
smaller space. The design of axial flow compressors is a great challenge, both
aerodynamically and mechanically.
The aerodynamic compressor design process basically
consists of mean line prediction calculation, through flow calculation, and
blading procedures. The mean line prediction is the first step within
compressor design. It is a simple one dimensional calculation of flow
parameters along the mid height line of the compressor where global parameters
as the annulus geometry, the number of stages, and the stage pressure ratios
are scaled 5. It is necessary to design axial flow compressor at preliminary
level and require parameters can be checked at initial level so further
improvement can be made at primary level before start a Detailed design.
It is a challenging job to design appropriate
compressor to meet the demands of VCE, which is the compressor should implement
performance adjustment of the engine and ensure the efficiency being maintained
within a higher range. As described by similarity principle, multiple
conditions can be converted to the same working condition according to the rule
of equality in reduced wheel speed, reduced mass flow and Mach number along
circumferential direction and so on. Then, the compressor performances under
different working conditions are approximately equal to each other. In this
research, a high loaded high-pressure compressor with high compression ratio
and large enthalpy rise was designed for VCE in two operating modes of low
bypass ratio (single bypass mode) and high bypass ratio (double bypass mode)
according to this principle 3.
complexities of a supersonic flow in compressor have been summarized: wave
structures such as expansive waves and compressive waves (even the shocks)
exist in supersonic regions. Due to this, flow parameter changes drastically in
the channel. Compressive waves may be formed by disturbed flow in that regions,
so influence on aerodynamic performances caused by compressive wave-boundary
layer interactions must be considered. The blade boundary layer, influenced by
blade’s geometrical parameters and main-flow aerodynamic parameters, usually
develops from laminar to turbulence. As a result, to capture the boundary layer
development, and estimating the transition position is significant for the
investigation of flow performances in the compressor. Besides, variations of
blade profile, blade stacking and end-wall effects often lead to 3D
characteristics in the flow fields, where secondary flows, separated flows and
complicated vortex structures exist. Indeed, it is important to obtain accurate
flow information and aerodynamic performances during the design process, which
is crucial for supersonic compressor.