# A Variable Cycle Engine (VCE) can be defined

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 1.

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

standby

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.

The 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.

The

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

condition.

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.

The

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.

1.

COMPRESSOR

DESIGN METHOD

The steps

involved in design of axial flow compressor is as shown in Figure 3 6:

By assuming values for the blade tip speed, and the

axial velocity and hub-tip ratio at inlet to the first stage, the required

annulus area at entry is obtained from the specified mass flow, assumed axial

velocity, and ambient conditions, using the continuity equation.

To

satisfy continuity equation:

From

the above equation,

Where

is considered as hub to tip ratio and density

is calculated from the relation

……..

(3) …ficient in both

modesgn phase of compressor.ip of rotor and stator blades.peed()be determined.

Thes

From the suitable design point under sea level conditions

(atmospheric pressure and temperature) and from cycle calculations, the

required pressure ration, mass flow rate and compressor inlet temperature can

be determined. These can be useful to calculate the stagnation and static

pressure and temperature at the first stage of compressor.

For the assumed axial velocity, tip radius will be the function

of hub to rip ratio.

Blade

speed can be represent as:

……. (4)

……. (5)

Therefore, we can iteratively determine the tip radius by

appropriate hub to tip ratio and rotational speed (n). It is useful to know

the mean radius as,

In

mean line design methodology, mean radius remain constant for all stages.

In case of exit area:

Blade height:

The radius at exit (final stage) of compressor as

determined as:

……. (9)

…….. (10)

The number of stages is found by dividing total temperature

rise in all stages by temperature rise per stage.

The

temperature rise for a stage is:

The

whirl components of velocity are determined from the velocity triangles as

shown in Figure 4.

The chord length of the blade airfoil is calculated

by appropriate selection of aspect ratio based on the application.

The

pitch and number of blades for the rotor are determined by the given relations,

……

(12)

4.1 THE APPLICATION OF SIMILARITY

PRINCIPLE:

The principle of

similarity is a consequence of nature for any physical phenomenon. By making

use of this principle, it becomes possible to predict the performance of one

machine from the results of tests on a geometrically similar machine, and also

to predict the performance of the same machine under conditions different from

the test conditions. If two machines are kinematically similar, the velocity

vector diagrams at inlet and outlet of the rotor of one machine must be similar

to those of the other. Geometrical similarity of the inlet and outlet velocity

diagrams is, therefore, a necessary condition for dynamic similarity 10.

VCE can combine the advantages of turbojet and

turbofan by adjusting its bypass ratio in order to change the Thermodynamic

cycle parameters. When the HP compressor was design, two working conditions

have been taken into account, which one is single bypass mode and the other is

double bypass mode. It means that HP compressor must adapt the bypass ratio

variations which makes the flow rate and compression ratio vary within wider

range than ones of conventional compressor besides the compressor must maintain

high efficiency. Namely, the flow fields in compressor should be reasonable

when the cycle parameters are changed greatly 7.

The outer bypass compress air flows into the inner

bypass when VCE changes its working conditions. Then the mass flow of HP compressor

increases. The change rate of mass flow depends upon the change of the bypass

ratio. There are some different working conditions for VCE matching with

different bypass ratios.

Assume that the

VCE works in n conditions. Then HP compressor must correspond with n series

of parameters as follows:

Condition 1

Condition 2

…………..

Condition i

…………..

Condition n

Obviously it is different from the conventional

compressor which is often given only one set of design parameters. It is very

difficult to design the compressor to meet so many series of cycle parameters.

For

turbo machinery, the total-total efficiency is commonly depended on ten

parameters, namely,

…… (14)

The

geometry and gas parameters can be omitted for the same compressor and the same

working fluid. The following equations can be gotten from the similarity

principle in turbo machinery.

……… (15)

Where,

Compression ratio

Reduced mass flow rate

Reynolds number

Reduced power

Reduced wheel speed

Both and are the qualitative

parameters and can be represented into:

…… (16)

…… (17)

Thus,

the efficiency is also represented into:

…… (18)

The effect of can be ignored if Re is

bigger than the second critical Re number in turbo machinery. Equation (18) is

reduced as:

…………… (19)

Therefore, if we make following equations both (20)

and (21) to be established for HP compressor of VCE, the multi-conditions

compressor for VCE can be treated as the conventional compressor to be

designed.

As a result, the mass flow rate and inlet parameters

of HP compressor are changed into any kind of condition among n series

of conditions when VCE adjusts its bypass ratio but the variations cannot influence

the efficiency of compressor, which is only influenced by the compression

ratio.

2. ANALYSIS OF DESIGN USING CFD TOOL

ANSYS Fluent 16.0, a CFD tool has been employed

within the computations to simulate the 3D steady flows and to validate the

aerodynamic performances of designed HP compressor for VCE. Both operating

conditions of single bypass mode and double bypass mode have been taken into

account.

Analysis

of design involves following steps:

·

Creating a geometry/mesh

·

Defining the physics of the model

·

Solving the CFD problem

·

Visualizing the results in post

processor

5.1 CREATING A GEOMETRY/MESH

The traditional approach to axial-flow compressor

aerodynamic design was to use various families of airfoil as the basis for

blade design. American practice was based on various families designed by the

National Advisory Committee for Aeronautics (NACA), the most popular being the

65-series family 8. NACA 65410 9 Airfoil is used here to generate blade

coordinates as in Figure 5.

To create the geometry NX UNIGRAPHICS 9.0 software

is used. NACA 65410 airfoil coordinates are imported to software and this

airfoil is used from entire hub to tip of rotor and stator blades. Meshing (Figure

6) has been carried out by Hyper Mesh. The type of element selected here is

Tetrahedral.

5.2 DEFINING THE PHYSICS OF THE MODEL

In this step we are defining the physics of Model.

This includes specification of type of fluid, defining the domains, Inlet and

Outlet Boundary conditions, type of analysis, turbulence model, and heat

transfer model etc. The following assumptions are taken for defining physics.

·

Steady state condition

·

No leakage losses

·

Friction between walls and fluid is

neglected.

5.3 SOLVING THE CFD PROBLEM

The

solver parameters are specified as follows:

·

Air as an Ideal gas is taken as Working

fluid.

·

The one-equation turbulence model of

Spalart-Allmaras has been applied to solve Reynolds’s averaged N-S equations,

and this model is applicable to the separation flow simulations of viscous

fluid under high pressure gradient.

·

For the two operating modes, total

temperature, total pressure and turbulent viscosity are prescribed at inlet;

the gas enters the calculation domain in axial direction. Meanwhile, static

pressure is given at outlet as shown in Figure 7.

·

Two domain interfaces are used. Rotor

domain is rotating and stator domain is stationary (Figure 7).

In

solver manger the number of iterations and accuracy is specified and results can

be analysed with help of CFD Post.

5.4 RESULTS OF CFD ANALYSIS

·

Static pressure distribution on blades

in two modes

·

Mach number distribution on the blades

·

Vortex movement

·

Pressure coefficient distribution on

blades in two modes.

6

RESULT

ANALYSIS

The results show that the relative errors of mass

flow compared to design value. In generally, the results of simulation indicate

that satisfactory aerodynamic performances are available in both of the two

operating conditions, and the designed HP compressor has achieved the anticipated

targets.

From the distribution of static pressure on blade

surface (not shown here), the fluid flows around the vane blade characterized

by pressure fluctuating along the whole blade height, and it is the most severe

near the hub, which is relevant to wave structures in the channel and

compressive wave boundary layer interactions. These influences are inevitable

in the supersonic or transonic compressor. In the stator, static pressure is

fully developed on the pressure surface but in the rotor, static pressure on

the blade presents certain differences in the two modes. It shows the power

requirement variation in two modes of operation.

Isolines distributions of absolute Mach number at

stator outlet in each mode represents the characteristics of supersonic wake

flow as depicted in Figure 8. The high-temperature gas flows through the

channel characterized as uneven at different blade heights because of blade

stacking and end wall effect. Mixing flow of supersonic fluid leads more energy

loss. Moreover, the uneven of flow parameters in the flow fields causes radial

movement from blade bottom to the top and corner vortex (Figure 9) near both

hub and shroud at outlet with secondary flow loss increasing apparently. These

secondary flow is mainly responsible for the aerodynamic loss in a supersonic

compressor.

Tip clearance flow is to blame for the aerodynamic

loss in both rotor and stator stage, with the proportion to be 45% and 30% in

some working conditions. It is significant to understand the features of

clearance flow for the design of rotor and to minimize the flow loss, so

representation formation of tip clearance flow in the supersonic compressor is analysed

with the aid of 3D streamlines as presented in Figures 10a-10c.

There are three forms of tip vortex, (1) Vortex

forms behind the guide vane, (2) Vortex forms due to the gas entering the top

gap from suction surface relative to high-speed rotating blades, (3) Vortex

formed by pressure gradient from blade pressure surface to suction surface

because of more tip clearance. Among these three, mass flow rate of the third

type of clearance flow is larger than the two previous kinds, and corner vortex

occurs when air current flows out of the suction side. So it is necessary to

control this vortex in design phase of compressor.

Wave structures of supersonic compressor are

analysed by the variation of absolute Mach number and pressure coefficient at

one particular blade height (Figure 11). Detached

shock forms when the gas flow out the trailing edge along pressure surface.

Intersection on different side occurs among detached shock, primary expansive

waves and reflective expansive waves. By the variations of Mach number and

pressure coefficient, it can be drawn the conclusion that the detached shock is

reduced and dissipated by the set of expansive waves. So, influence of the

detached shock is limited to the flow fields near the trailing edge, while the

expansive wave can occupy most space in the tapered-cut

region.

(b) Double

bypass mode

Figure 11: Mach number

and pressure coefficient in both modes

7

CONCLUSION

In present paper, a

high loaded high-pressure compressor stage with high compression ratio has been

designed by NACA profile based on similarity principle. Three-dimensional

calculations have been carried out under different thermodynamic cycle

parameters to simulate flow fields in the two operating conditions (Single

bypass mode and double bypass mode). It reveals the characteristics of supersonic

wake flow in a