Gas Turbine Engines: Turbojet, Turbofan & Turboprop

Inside the jet engine. Learn the Brayton cycle, engine stations (inlet, fan, compressor, combustor, turbine, nozzle), spool configurations, bypass ratio, compressor and turbine aerodynamics, surge and stall margin, combustor design and emissions, turbine cooling and materials, afterburners, turboprops and turboshafts, engine performance and FADEC control — with Python for ideal cycle efficiency and specific work.

A gas turbine squeezes air (compressor), burns fuel in it (combustor), extracts enough energy to drive the compressor (turbine), and uses the rest for thrust (jet nozzle, fan) or shaft power (propeller, helicopter rotor, generator).

Modern turbofans are among the most sophisticated machines ever built: turbine blades run in gas hotter than their own melting point, spinning at thousands of rpm, for tens of thousands of hours between overhauls.

An analogy

'Suck, squeeze, bang, blow': the inlet sucks air in, the compressor squeezes it, the combustor bangs (burns) fuel, and the turbine/nozzle blows it out — the same four strokes as a car engine, but happening continuously.

Engine stations

Inlet (diffuses air to ~M 0.5), fan (in turbofans), low- and high-pressure compressors (axial stages of rotors and stators), combustor (fuel burns at ~constant pressure), high- and low-pressure turbines (drive the compressors and fan through concentric shafts/spools), and the nozzle (accelerates the exhaust).

Engine types

Same core, different use of its energy:

Compressors, surge and control

Axial compressors add a little pressure per stage (~1.3–1.5) through many stages; each blade row is an airfoil that can stall. If too many stall, the whole compressor surges (flow reversal, bang, flame out the inlet). Variable stator vanes, bleed valves and multiple spools keep a stall margin. The FADEC (full-authority digital engine control) schedules fuel flow, vanes and bleeds to deliver commanded thrust while protecting limits (N1, N2, EGT, surge).

Hot section, emissions and the future

Turbine inlet temperatures of 1700–2000 K exceed nickel superalloy melting points; blades survive through single-crystal alloys, internal cooling passages with film cooling holes, and thermal barrier coatings (ceramic matrix composites are the next step). Lean-burn combustors reduce NOₓ. Future directions: higher OPR and BPR, sustainable aviation fuels, hydrogen combustion, hybrid-electric architectures, and open rotors.

Formulas

Worked example: Ideal efficiency

Given: OPR = 40

Find: η_th.

Answer: 65.1%

Worked example: Compressor exit temperature

Given: T₁ = 288 K, π = 20, η_c = 0.88

Find: T₂.

Answer: ≈ 731 K

Worked example: Bypass thrust share

Given: BPR 10, core V_e = 450, bypass V_e = 300, V₀ = 240 m/s

Find: Bypass share of thrust.

Answer: 600/810 ≈ 74% from the fan

Ideal Brayton cycle

# Ideal Brayton cycle / turbojet core
cp, g, T1 = 1005.0, 1.4, 288.0
for pr in (10, 20, 40):
    eta = 1 - pr ** (-(g - 1) / g)
    print(f"pressure ratio {pr:2d}: ideal thermal efficiency {eta*100:.1f}%")
# Specific work for T3 = 1600 K, pr = 20
pr, T3 = 20, 1600.0
T2 = T1 * pr ** ((g - 1) / g)
T4 = T3 / pr ** ((g - 1) / g)
w_net = cp * ((T3 - T4) - (T2 - T1))
print(f"T2={T2:.0f} K, T4={T4:.0f} K, net work {w_net/1000:.0f} kJ/kg")
# pressure ratio 10: ideal thermal efficiency 48.2%
# pressure ratio 20: ideal thermal efficiency 57.5%
# pressure ratio 40: ideal thermal efficiency 65.1%
# T2=678 K, T4=680 K, net work 533 kJ/kg

Efficiency rises with pressure ratio (modern OPR 40–60); specific work (thrust per kg of air) rises with turbine inlet temperature. Real engines lose several points to component inefficiencies, cooling air and pressure losses — hence the relentless push on turbine materials and cooling.

Common mistakes

Applications

Key takeaways

Going further