Engineering Design of Recuperative Gas Combustion Chamber for Gas Turbine Engine

1. • Apurva Anand (1NT13AE009)
• Prabin Sherpaili (1NT13AE037)
• Roshan Sah (1NT13AE052)
• Ghanendra Kumar Das (1NT13AE067)
Under Guidance of
• Dr N K S Rajan, Chief Research Scientist,
Department of Aerospace Engineering
Combustion-Gasification-Propulsion Lab, IISC
• Asst Prof. Harish H V,
Department of Aeronautical Engineering
NMIT

2. Objective
 To design a can-type combustion chamber for producer gas fuel for commercial
use
 To develop passive heat recovery technique in order to minimise conduction
losses in the combustion chamber
 To develop engineering model and drafts for production of the combustion
chamber for testing at CGPL lab, IISc.

3. 1. The length of the combustion chamber must be less than one metre.
2. The diameter of the combustion chamber must be less than 500 mm.
3. The heat recovery technique must be passive and must ensure 200K-300K
rise in exhaust temperature.
4. The maximum pressure difference between any two points should be less
than 150 mmWC
Performance Requirements

4. • Use of Counter flow recuperative heat exchanger for gas combustion chamber.
 Enhanced performance but least additional space
 Suitable to be used in Gas turbines
• Use of cross pattern fins
 Increases heat transfer coefficient
 Induces swirling and turbulence for proper mixing
Design Exit temperature
Without recuperation 834 K
With recuperation 961 K 15%
increase
With recuperation
and fins
1200 K 44%
increase

5. mfxCv = maxCpax(Tfinal – Tinitial of air)+mfxCpfx(Tfinal–Tinitial of fuel )
mf α
𝟏
Tinitial of air
for constant output temperature required
and
Tfinal α Tinitial of air for same mass of fuel consumed
INCREASED EFFICIENCY

6. air at initial temp
air inside preheat section
Inner Combustion Chamber
Fig 1: Counter Flow Recuperation
Cold air inlet

7. Equations for theoretical estimation
• Combustion Equation
0.2 CO +0.2 H2 +0.01 CH4 +0.11 CO2 +0.48 N2 +x(O2 +3.76 N2)-- 0.32 CO2 +0.48 N2 +0.22
H2O +3.76x N2
• Reynold no of internal flow in a cylinder (Re) = ƍ*u*D1/ µ
• Nusselt Number, Nu=0.023*Re^0.8* Pr^0.4
• Heat transfer coefficient (h1) = Nu* K/D1
• Logarithmic mean temperature difference (ɵm) =
ɵ2−ɵ1
𝑙𝑛
ɵ2
ɵ1

8. air
air
wall
Fig: Electrical Analogy
Counter flow heat exchanger
Th1=1060KTc1=400K
Th2=977KTc2=300K
Section Ti(K) To(K) Hi(W/m2
K)
Ho(W/m
2K)
Wall
temp(T0’)
Hot section 1060 400 22.2 30.7 642.26
Intermediate
section
1018 350 21.92 34.266 576.75
Cold Section 977 300 21.6 36.25 520
L=0.86187m
• Final Inlet air
temperature=401K
• CONDUCTION LOSS=
11%
• Heat recovered=8%

9. 633 K
(642K expected)
Verification of Numerical Simulation

10. Product development methodology

11. Particular Detail
Mass flow rate 0.028517 kg/s
Fuel flow rate 5.0893e-3 kg/s
Air Fuel Ratio 5.6
Length of combustor 715 mm
Inner diameter of Can 100 mm
Hydraulic Diameter of recuperative passage 20 mm
Fuel inlet diameter(each) 9 mm
No of Fuel Supply inlets 8
Velocity of fuel supply at fuel inlets 10 m/s
Velocity of air in recuperative passage 6.5 m/s
No of fins 100
Dimension of fins 85mmx2mm at 20 degree with axis
alternatively
Material Steel
Total Mass of the combustion chamber 21 kg
Total Volume 3 litres
Estimated Cost of Fabrication INR 30,000-35,000

13. Structured Meshing in ICEM
Unstructured Meshing in Ansys Workbench

14. INPUT BOUNDARY CONDITIONS
Solver: Fluent
Inlet: Velocity Inlet
Exit: Pressure exit
Modelling: K-epsilon model
Combustion type: Non-premixed combustion
Rich Flammability Limit: 0.7
Component Air (%) Producer Gas (%) Air (%)
Carbon dioxide 0.11 0.03
Carbon Monoxide 0.20 0
Hydrogen 0.20 0
Methane 0.01 0
Nitrogen 0.48 0.48
Oxygen 0 0.21

15. Fig: Fuel Flow Streamlines
Fig: Mixture Fraction Distribution

16. Particular Detail
Maximum exit temperature 860 K
Temperature of Air at inlet after heat exchange 300 K
Maximum Pressure drop 49. mm WC
Combustion Without any Recuperative technique

17. Particular Detail
Maximum exit temperature 1200 K
Temperature of Air at inlet
after heat exchange
560 K
Maximum Pressure drop 100 mm WC
Fig: Temperature plane at exit
Fig: Temperature plane at inlet to combustion chamber

18. Fig: Temperature at
exit=1200 k
Fig:Max Pressure
Difference= 100
mmWC
Fig:Swirling
action due to
fins

19. Component Mole percentage
CO2 32%
N2 48%
H2O 20%

20. CONCLUSION