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Design of Transonic Axial Compressor

Sai Kiran Goud.M

Transonic axial flow compressors are today widely used in aircraft

engines to obtain maximum pressure ratios per single-stage.

. High stage pressure ratios are important because they make it

possible to reduce the engine weight and size and, therefore,

investment and operational costs.

Performance of transonic compressors has today reached a high

level but engine manufacturers are oriented towards increasing it

further.

Another important target is the improvement of rotor stability

towards near stall conditions, resulting in a wider working range.

Today’s high efficiency transonic axial flow compressors give a total

pressure ratio in the order of 1.7-1.8.

Introduction

Based on multistage axial flow compressor of the M7A-O2 gas

turbine establishes a higher pressure ratio and a large amount of

air flow than the M7A-O1.

By adopting transonic stages with the characteristics of high

pressure ratio, the air flow rate and pressure ratio are increased.

A preliminary design for a nine-stage axial flow compressor was

produced from an initial specification. This design is presented

here, and is criticized. Alterations are then made to the design to

compensate for its deficiencies.

Problem Definition

By considering the initial required conditions for the design of axial compressor the following parameters are designed.

All the blade angles at mean for both rotor and stator of the compressor

Total pressure ratio of each stage and the complete compressor.

Total shaft power.

Assuming every stage has the same mean radius with rotor blade angle changing by ‘12 degree’

Number of stages n 9

Rotating Speed N 9000rpm

Ambient pressure P01 15.7Psia

Ambient temperature T01 519R

Mean axial velocity Va 520ft/s

Radius of the rotor at tip rt 12inches

Radius of the rotor at hub rh 8inches

Change in rotor blade angle per

stage

β1 – β2 12degree

Design Specifications

Initial design specification

Calculate the mean radius rm= rt+ rh/2 = 10. Therefore, hub/tip ratio = 0.83ft.

Calculate the blade velocity ‘Um’ at the mean, by using the formula

Um= 2π.N.rm / 60 = 781.86ft/s Calculate the inlet blade angle β1 ,

Tan β1 = U / V1 = 56.30 degrees As given, Change in rotor blade angle per stage, β1 –

β2= 10, calculate β2 = 46.30 degrees.

From the velocity triangle , find the value of Wu2, Tan β2 = Va / Wu2 = 496.9 ft/s

Procedure / Calculations

Velocity Triangle

In velocity triangle we can get the value of Vu2 by subtracting it from ‘U’ = 284.96ft/s

Using velocity triangle and the value of Wu2 , calculate ‘α2’ , i.e.

Tan α2 = Vu2 / Va = 28.36degree Now find stagnation enthalpy difference

by using the formula,Ϫ hos= U. Vu1 = U (U- Va tan β2) =

7.42btu/lbm

By using Cp (To2-To1) = Ϫ hos, find temperature per stage

T o2 = 539R

Pressure ratio per stage , as T o2 = T03, calculate P03/ P01,

P03/ P01 = ( T03 /T01)k/k-1 =1.145

Calculate the outlet temperature of the compressor, Toe

Cp (Toe-To1) = Ϫ hos = 703.5R

Calculate the overall pressure ratio of the compressor,

Toe / To1 = (Poe / Po1)k-1/k = 2.89

Using the above equation find out ‘Poe’. = 45.5psia

For finding the pressure rise per stage, calculate T1,

T1= T01- V12/ 2 Cp =496.51R

Now calculate P1,

P1/P01 = (T1/ T01)k/k-1 = 13.44psia

Now calculate 𝝆1, by using the equation,𝝆1= P1 / RT1 = = 0.0507

lbm/ft3

Calculate area of the rotor ,A1= π (rt

2- rh2)

=251.2ft2

Calculate the mass flow rate,M .= 𝜌1. V1. A1 = 66.57

lb/s Finally calculate the shaft power,

Ps= m..n. Ϫ hos = 4445.54hp

Inlet blade angle at mean β1 56.30degree

Outlet Blade angle at mean β2 46 degree

Blade angle at mean α2 28.36degree

Pressure ratio per stage P03 / P01 1.145

overall pressure ratio Poe / Po1 2.89

mass flow rate m' 66.57lb/s

shaft power Ps 4445.54hp

Results

Results

Hence, for a transonic axial compressor, determined the following values,

1. Blade angles at mean for both stator and rotor.

2. Total pressure ratio of each stage and the complete compressor.

3. Total shaft power.

Conclusion

Thank You