How Turbojet Engine Works?

     The turbojet is an air-breathing jet engine, usually used in aircraft. It consists of a gas turbine with a propelling nozzle. The gas turbine has an air inlet, a compressor, a combustion chamber, and a turbine (that drives the compressor). The compressed air from the compressor is heated by the fuel in the combustion chamber and then allowed to expand through the turbine. The turbine exhaust is then expanded in the propelling nozzle where it is accelerated to high speed to provide thrust. These modern engines use a gas turbine engine core with a high overall pressure ratio and high turbine entry temperature and provide a great deal of their thrust with a turbine-power fan stage, rather than with pure exhaust thrust as in a turbojet. These features combine to give a high efficiency relative to a turbojet. All practical air-breathing jet engines are internal combustion engines that directly heat the air by burning fuel, with the resultant hot gases used for propulsion via a propulsive nozzle.

     It mainly works on the principle of Brayton Cycle where various thermodynamic processes take place in stages. Ambient air is drawn into a compressor, where it is compressed; ideally an isentropic process. The compressed air then runs through a mixing chamber where fuel is added, an isobaric process. The pressurized air and fuel mixture is then ignited in an expansion cylinder and energy is released, causing the heated air and combustion products to expand through a turbine; another ideally isentropic process. Some of the work extracted is by the turbine to drive the compressor through a crankshaft. The rest of the power is used to propel the aircraft.

Figure 1. A 2-D image of a centrifugal type Turbojet Engine
The complete engine consists of many parts that are necessary for its effective functioning. But the following parts are the once which can be called as organs of a turbojet engine as shown in Figure 1.
  • Air Intake structure Compressor
  • Combustion Chamber
  • Turbine
  • Nozzle
Each of these components has their respective functions that are required for the running of the engine. These components and their functions have been explained in detail along with the design aspects later.

     The compressor is driven by the turbine. It rotates at high speed, adding energy to the airflow and at the same time squeezing (compressing) it into a smaller space. During the design following assumptions were made:

  • Centrifugal compressor was designed as it is easier and comparatively less expensive to manufacture and provides better efficiency at lower RPM. 
  • Calculations are done for 8000RPM. 
  • Inlet angle is 60°. 

The values like impeller dimensions were obtained by continuous trial and error iterations have done using Vista CCD, a tool in Ansys Workbench.

Formulae Used :
  • Tangential velocity at the inlet, π‘ˆ1= πœ‹∗𝑑1∗𝑁/60000 (1) 
  • Axial velocity at the inlet, 𝑉𝑓1= tan (𝛼)* π‘ˆ1 (2) 
  • Inlet area, 𝐴1= πœ‹∗ (𝑑12−𝑑22)/4 (3) 
  • Flow area, 𝐴2= πœ‹ ∗ 𝑑3∗ 𝑙 (4) 
  • Continuity equation, 𝐴1𝑉𝑓1= 𝐴2𝑉𝑓2 (5) 
  • Mass flow rate at outlet, π‘ṧ = 𝜌 ∗ 𝐴2∗ 𝑉𝑓2 (6)  
  • Temperature at outlet of compressor T2= T1*(p2/p1)(0.4/1.4) (7) 
  • Enthalpy Ξ”h1= CP*Ξ”T (8) 
  • Power required P= (m ̇*Ξ”h1)/Ξ·mech (9)

     With the above equations [4] the various parameters are like velocity, mass flow, cross-section areas, Temperature, enthalpy, and power are calculated. These results are used for the sketches and the Ansys work. The results of the various parameters are shown in Table 1. The sketch of the impeller with its 3 views is shown in figure 2.

     The pressure at the outlet of the impeller was set to 2.9 bar [5]. It is evident from figure 3, that the pressure developed by each blade is 2.9 bar. This is represented by the green region. The blade produces the pressure required at the outlet, hence the design is satisfactory.

 A combustor is a component or area of a gas turbine where combustion takes place. The objective of the combustor in a gas turbine is to add energy to the system to power the turbines and produce a high-velocity gas to exhaust through the nozzle in aircraft applications. The burning of fuel must be complete, otherwise, the engine is wasting the unburnt fuel and creating unwanted emissions of unburnt hydrocarbons, CO and soot. The turbine which the combustor feeds, need high-pressure flow to operate efficiently. The flame must be held inside the combustor. If the combustion takes place further back in the engine, the turbine stages can easily be damaged. Space and weight are at a premium in aircraft applications, so a well-designed combustor strives to be compact. There are strict regulations on aircraft emissions of pollutants like carbon dioxide and nitrogen oxides, so combustors need to be designed to minimize those emissions.

     In order to match the required design goals, the combustor primary zone is designed with a fuel-rich mixture at an equivalence ratio of 0.95. A primary zone airflow rate of 0.1684 kg/sec was found to perform good ignition performance, and promising low NOx emissions and high combustion efficiency at low power conditions with the expenses of increased exhaust smoke and reduced volumetric heat release rate [7]. The distribution of the primary zone air is such that 33.34% I.e.0.027 kg/sec enters through 52 holes of 2.2 mm diameter located in the dome section and the rest 66.6% i.e. 0.0531 kg/sec, flows in the annulus and enters through 102 holes with 2.2 mm diameter each and arranged in two rows of 51 holes each. The remaining airflow is used for both secondary and dilution zones as well as for film cooling of the liner at the outer and inner locations. In the secondary zone, 0.133 kg/sec of air was necessary to achieve an equivalence ratio of 1.7. The total number and size of the secondary zone air injection holes are found 304. Then these holes are in turn in a manner that every set contains 152 holes arranged in two rows of 76 holes of 2.2 mm diameter each.

     In the dilution zone, an amount of 0.367 kg/sec of air enters through 16 holes of 16 mm diameter each located on the outer liner. In order to protect the liner walls from overheating, it is necessary to introduce an amount of airflow that will form a barrier against the hot following gas. Accordingly, 6.58% of the chamber total airflow rate i.e. 0.053 kg/sec enters in the annular space between the inner liner and engine shaft to film cool the liner wall at this location and distributed through 96 holes total of 2.3 mm in diameter each. These holes are in turn arranged in two sets of 48 holes each that each set, in turn, contain two rows of 24 holes. This arrangement assures the proper air distribution and guarantees maximum protection to the inner liner wall. An amount of airflow rate is required to enter the annulus passage of the combustor to film cool the outer liner at this location and 9.53% of the chamber total airflow rate i.e. 0.079 kg/sec is admitted through 152 holes total with 2.2 mm diameter each. And in order to distribute this air properly to perform the required task of the liner wall protection from overheating, these holes are in turn arranged in two rows of 76 holes of 2.2 mm diameter each.

Figure 4. 2-D drawing of combustion chamber outer liner

Figure 5. 2-D drawing of combustion chamber inner liner

     The calculation of the combustion zone length is based on the chamber volume, inlet velocity, and combustion product's residence time. The combustion chamber volume is in turn determined on the basis of the stirred reactor model [7, 9]. The primary zone products residence time is found 1.317 ms, while the air velocity at the inlet to this zone is 32.5 m/sec. While the primary zone length is found 85.6 mm. The secondary zone length is determined by the procedure found in [7] i.e. on the basis of achieving just enough residence time to complete the reaction process and to consume the high levels of primary zone carbon monoxide and unburned fuel before the gas enters the dilution zone. And accordingly, the gas residence time in the secondary zone is found 1.003 ms, while the gas velocity at the inlet to this zone is 36.915 m/sec. Thus the secondary zone length is found to be 74 mm.

     The design process involves consideration of power and efficiency, as well as weight, cost, volume, life, noise, etc. The first major step in design is to carry out the thermodynamic design point calculations. These include important factors such as component efficiencies, air bleeds, variable fluid properties, and pressure losses. The choice of cycle parameters is strongly affected by the engine size, especially the air mass flow rate. Small engines have small blades, which cannot readily be cooled (the manufacturing complexity and cost would be excessive). 

     Hence the pressure ratio may be restricted to allow the blading to be of the airflow, pressure ratio and Turbine Inlet Temperature (TIT) is known, and attention can be turned to the aerodynamic design of the turbomachine, to determine annulus dimensions, rotational speeds and the number of stages. In this work, the detailed aerothermal design of a multi-stage axial flow turbine with air cooling is presented, with consideration of the particulars of blade and nozzle geometry, the limiting conditions on the Mach number, annulus divergence and number of blades.

Formulae Used :
  • Tangential speed, U= (Ο€*D*N)/60 (10) 
  • Turbine exit blade angle tanΞ²3 = tanΞ±3 +1/Ξ¦ (11) 
  • Temperature change across the stage  = (2*Cp*Ξ”Tos)/U2 (12) 
  • Stage reaction TanΞ²3= (½)*∅ (Ο†/2+2R) (13) 
  • Turbine inlet flow angle tanΞ±2= tan (Ξ²2+ 1/∅) (14) 
  • Pressure ratio across stage Ξ”Tos= Ξ·*T01 [1-(p03/p01) ((Ξ³-1)/Ξ³)] (15)
  • Absolute velocity C2= Ca/cos (Ξ±2) (16) 
  • Temperature equivalent of kinetic energy T02-T2= (C22) / (2*Cp) (17) 
  • Pressure at turbine inlet (p01/p2) = (T01/T2) (Ξ³/ (Ξ³-1)) (18) 
  • Density at turbine inlet ρ2=p2/(R*T2) (19) 
  • Area at Turbine inlet A2= m ̇/(ρ2*Ca2) (20) 
  • Temperature at turbine exit T03= T01-Ξ”Tos (21) 
  • Density at turbine outlet ρ3= p3/(R*T3 ) (22) 
  • Area at turbine outlet A3= m ̇/ (ρ3*Ca3) (23) 
  • Density at nozzle inlet ρ1= p1/ (R*T1) (24) 
  • Area at nozzle inlet A1= m ̇/ (ρ1*Ca1) (25) 
  • Average area A avg = h*(2*Ο€*rm-t*n) (26)

Figure 6. 2-D drawing of Stator

Figure 7. 2-D drawing of Rotor

With the above equations [4] the various parameters are like velocity, speed, blade angles, mass flow, cross-section areas, Temperature, enthalpy, density, power, etc are calculated. These results are used for the sketches and the Ansys work. The results of the various parameters are shown in Table 4. The sketch of the Stator and the Rotor are shown in figure 5 and figure 6.

     A nozzle is a device designed to control the direction or characteristics of fluid flow. In a turbojet engine, the main function of a nozzle is to increase the outlet velocity of the exhaust gases so as to achieve maximum output thrust which is dependent on output characteristics of the exhaust gases. A convergent nozzle has been designed considering that there shall be no surging or reverse flow of gases and the critical area was calculated using an equation.
" Ac= m ̇ / (216.5*(p01/v01)0.5) (27) "
     At critical area the velocity of the fluid will be sonic and are lower than that will lead it to supersonic speed and needs divergent section but if the exit is higher than that of critical area divergent section is not needed and there will not be any choking. Considering,
" de = 0.1 m (28) "
     With the above equations [4] the various parameters are like velocity, mass flow, cross-section areas, Temperature, density, power are calculated. The results of the various parameters are shown in Table 5.

Powered by Blogger