Pre-Stall Behavior of a Transonic Axial Compressor Stage ...
Design of transonic axial compressor
Transcript of Design of transonic axial compressor
Design of Transonic Axial Compressor
Sai Kiran Goud.M
Introduction
• 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.
Problem Definition
• 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.
• 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’
Design Specifications
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
Initial design specification
Procedure / Calculations
• 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
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
Results
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
Conclusion
• 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.
Thank You