Gas turbines and jet propulsion

In short

A gas turbine burns fuel in a continuous flow of compressed air and expands the hot gas through a turbine; part of the turbine work drives the compressor and the rest appears as shaft power or as a high-velocity jet. The same machine therefore serves two families of application — power generation and mechanical drive on the ground, and aircraft, missile and marine propulsion. It is examined heavily in GATE (ME and AE), ESE, ISRO/DRDO and PSU papers, usually as actual-cycle calculations with component efficiencies, regeneration, thrust and propulsive efficiency. In India the knowledge is used at NTPC's, GAIL's and ONGC's gas-based and combined-cycle stations, at BHEL and Siemens Energy, at HAL's Engine Division and GTRE (Kaveri), at ISRO, and at the GE Aerospace, Safran, Rolls-Royce and Honeywell engineering centres in Bengaluru, Pune and Hyderabad.

Explained

The gas turbine plant and its classification

• Core components: air intake and filter, compressor (axial, multistage, in large units; centrifugal in small engines), combustion chamber, turbine, exhaust diffuser or propelling nozzle, plus starter, fuel system and controls.
• Open cycle: atmospheric air is taken in, burnt with fuel and exhausted — the working fluid changes, combustion is internal. Almost all practical gas turbines, aero and industrial, are open cycle.
• Closed cycle: a fixed charge of gas (helium, CO₂, air) circulates; heat is added in a heater (external furnace or nuclear reactor) and rejected in a pre-cooler. Advantages: any fuel, clean turbine blades, higher density so smaller machine, output controllable by changing pressure level. Disadvantages: bulky heat exchangers, heater metal temperature limits the cycle, low power-to-weight — so it is never used in aircraft. Semi-closed cycles mix the two.
• Shaft arrangements: single-shaft (compressor, turbine and load on one shaft — good for constant-speed generator drive); twin-shaft / free power turbine (gas generator spool plus a separate power turbine — better part-load and starting torque, used in mechanical-drive and marine sets); multi-spool aero engines (LP and HP spools on concentric shafts).
• Constant-pressure (Brayton/Joule) combustion is universal; the constant-volume (explosion) gas turbine is only of historical interest.
• Why gas turbines are attractive: very high power-to-weight and power-to-volume ratio, no reciprocating parts so low vibration, few minutes to full load (so ideal for peaking and grid balancing), low lubricating-oil consumption, can burn natural gas, naphtha, HSD or kerosene.
• Limits: low part-load efficiency, output falls sharply with ambient temperature and altitude, high air–fuel ratio (about 50:1 to 100:1) needed to protect the blades, costly high-temperature materials, noisy, and a large share of turbine work is eaten by the compressor.

From the ideal cycle to the actual gas turbine

• Ideal air-standard efficiency depends only on pressure ratio: η = 1 − 1/t where t = rp^(γ−1)/γ. Raising rp raises efficiency but net work peaks and then falls, because compressor work grows faster than turbine work.
• Real losses: irreversible compression and expansion (isentropic efficiencies ηc ≈ 0.85–0.90, ηt ≈ 0.88–0.92), 2–4% combustor pressure loss, 1–2% intake and exhaust duct loss, 98–99% combustion efficiency, turbine cooling-air bleed (up to 15–20% of core air), mechanical and bearing losses, and variable cp and γ with temperature (gases: cp ≈ 1.15 kJ/kg·K, γ ≈ 1.33).
• Work ratio = net work ÷ turbine work. A gas turbine's work ratio is only about 0.3–0.5 (a steam plant's is above 0.99), so a small drop in ηc or ηt causes a large drop in net output — this is the single most important practical fact about gas turbines.
• Turbine inlet temperature (TIT) is the other governing variable: for a given pressure ratio, raising TIT raises both net work and efficiency. Modern heavy-duty F/H-class machines run about 1,400–1,600 °C firing temperature; aero engines run higher still, above the melting point of the blade alloy, which is why cooling is essential.
• Effect of ambient conditions: industrial sets are rated at ISO conditions — 15 °C, 1.013 bar, 60% RH. Output falls roughly 0.5–0.9% per °C of ambient rise and with altitude, so an Indian summer site can lose 15–20% of nameplate output. Remedies: inlet evaporative cooling or fogging, inlet chilling, and online/offline compressor water washing to remove fouling.
• Compressor aerodynamics: stage pressure ratio 1.15–1.4, so 10–20 stages give rp of 10–40. Surge and rotating stall are avoided with variable inlet guide vanes, variable stators and bleed valves; the surge line on the compressor map limits acceleration.
• Governing: fuel flow control for load, variable geometry, and in generator service a droop/isochronous speed governor; aero engines schedule fuel to keep spool speed and exhaust gas temperature within limits.

Improving the simple cycle

• Regeneration (heat exchange): turbine exhaust preheats the compressed air before the combustor, cutting fuel. It only works while turbine exit temperature exceeds compressor exit temperature, i.e. at low pressure ratios — so regenerative plants use modest rp. Effectiveness ε = heat actually recovered ÷ maximum possible, typically 0.70–0.90. Regeneration raises efficiency but not net work.
• Intercooling: compression split into LP and HP with a cooler between; reduces compressor work and increases net work, but lowers the air temperature entering the combustor so more fuel is needed — efficiency falls unless a regenerator is added.
• Reheat: expansion split into HP and LP turbines with a second combustor between; increases turbine work and net work, raises exhaust temperature (useful for regeneration or a waste-heat boiler) but by itself lowers efficiency.
• Intercooled–reheat–regenerative (ICR) cycle: all three together give the best combination of high specific work and high efficiency, at the cost of plant complexity — used in some marine and aeroderivative designs.
• Combined cycle (CCGT): gas turbine exhaust at 500–650 °C feeds a heat recovery steam generator (HRSG) driving a steam turbine. Net plant efficiency about 52–60% (LHV), the highest of any thermal power plant. Typical split: two-thirds of output from the gas turbine, one-third from the steam turbine.
• Cogeneration / CHP: exhaust heat used as process steam or for absorption chilling; overall fuel utilisation can exceed 80%. Common in Indian fertiliser, refinery, textile and paper plants.
• Other developments: steam or water injection (STIG) for power boost and NOx control, dry low-NOx (DLN) premixed combustors, IGCC with coal gasification, and supercritical CO₂ closed cycles under research.

Combustion chamber, blades and materials

• Combustor types: tubular (can) — several separate flame tubes, easy to maintain, used in older industrial sets; annular — single ring, compact, lightest, used in most modern aero engines; can-annular (tubo-annular) — compromise used in heavy-duty machines.
• Zones: the primary zone behind a swirler burns near-stoichiometric at high temperature; the secondary (intermediate) zone completes burn-out; the dilution zone mixes in the remaining air to bring the gas down to the allowable turbine inlet temperature with an acceptable radial temperature profile (pattern factor).
• Requirements: high combustion efficiency, low pressure loss, stable flame over a wide range, wide ignition and stability limits, no carbon or smoke, low NOx, and reliable relight at altitude.
• Blade cooling: internal convection cooling through serpentine passages, impingement cooling on the leading edge, film cooling through surface holes, and transpiration cooling; plus thermal barrier coatings (yttria-stabilised zirconia) and anti-oxidation coatings.
• Materials: titanium alloys and steels for compressor blades and discs; nickel-base superalloys (Inconel, Nimonic) for turbine blades, made as equiaxed, directionally solidified or single-crystal castings to resist creep; ceramic matrix composites are entering hot sections.
• Failure modes to know: creep, thermal fatigue, hot corrosion and sulphidation (from fuel sulphur and sodium/vanadium), foreign object damage, blade tip rub, and compressor fouling.

Jet propulsion: thrust and the efficiencies

• A propulsion engine works by Newton's third law: it accelerates a mass of fluid rearwards, and the reaction is thrust. Momentum thrust comes from the change of momentum; pressure thrust appears only when the nozzle is not fully expanded.
• Propulsive (Froude) efficiency = useful thrust power ÷ (thrust power + kinetic energy wasted in the jet) = 2/(1 + Cj/Ca). It rises as the jet velocity approaches flight speed — so for a given thrust it is better to give a large mass of air a small acceleration (propeller, high-bypass fan) than a small mass a large acceleration (turbojet).
• Thermal (internal) efficiency = rate of kinetic energy added to the fluid ÷ fuel energy supplied; overall efficiency ηo = ηp × ηth = F·Ca ÷ (ṁf × CV). Note that at zero flight speed propulsive and overall efficiency are zero even though thrust is maximum.
• Performance measures: specific thrust F/ṁa (N·s/kg) — determines engine frontal area; TSFC (thrust specific fuel consumption, kg/N·h) — determines fuel burn; and for rockets specific impulse Isp = F/(ṁ g₀) in seconds.
• Nozzles: convergent nozzles are normal and choke at a pressure ratio of about 1.89, with the residual expansion giving pressure thrust; convergent–divergent nozzles are used for supersonic aircraft and rockets. Thrust reversers (clamshell, cascade, cold-stream) shorten landing runs.
• Thrust augmentation: an afterburner (reheat) injects fuel into the jet pipe downstream of the turbine, raising thrust by 40–70% with a two- to three-fold rise in TSFC; water or water–methanol injection was used for take-off boost on older engines.
• Thrust lapse: thrust falls with altitude roughly in proportion to air density, and (for a turbojet) changes only mildly with speed because ram pressure rise offsets the rising intake momentum drag.

Types of propulsion engines and where each is best

• Turbojet: all thrust from the core jet. Simple, small frontal area, best above about Mach 1.5–2 (with afterburner on fighters), poor fuel economy and very noisy at low speed.
• Turbofan: a fan driven by extra turbine stages passes bypass air around the core. Bypass ratio (BPR) = bypass flow ÷ core flow; low BPR (0.3–1) with afterburner for combat aircraft, high BPR (5–12) for airliners. High BPR gives better propulsive efficiency, lower TSFC and much lower noise at Mach 0.8–0.85 — the airline standard (CFM LEAP, PW GTF on Indian A320neo fleets).
• Turboprop: most of the gas energy is taken by the turbine and geared down to a propeller; about 85–90% of thrust from the propeller, the rest from the jet. Best economy below about 700 km/h and on short sectors — ATR 72 and the HAL-built Dornier 228 in Indian regional service.
• Turboshaft: all output as shaft power — helicopters (HAL Dhruv, LCH), marine propulsion, and industrial/aeroderivative gensets and pipeline compressors.
• Ramjet: no compressor or turbine; the intake diffuser compresses by ram effect. Cannot start from rest, needs boosting to about Mach 0.5–1, efficient at Mach 2–4. Used in missiles; the scramjet burns in a supersonic stream above about Mach 5 (ISRO and DRDO have flight-tested scramjet and air-breathing technology demonstrators).
• Pulsejet: intermittent combustion behind flap valves — crude, noisy, now only in targets and hobby engines.
• Rocket: carries both fuel and oxidiser, so thrust is independent of flight speed and it works in vacuum; highest thrust per unit weight, very high propellant consumption. Solid (PSLV strap-ons, S200 of LVM3), liquid (Vikas engine, CE-20 cryogenic upper stage) and hybrid types.

Indian standards, codes and practice

• Performance and acceptance testing: ISO 2314 (gas turbines — acceptance tests), ISO 3977 series (procurement), ASME PTC 22 (gas turbine performance test), ASME PTC 46 for overall plant; API 616 for gas turbines in petroleum and gas service; vibration to ISO 10816/20816.
• Power sector: gas-based generation is licensed and tariffed under the Electricity Act, 2003 with CERC/SERC regulation and CEA technical standards; gas-based capacity is roughly 25 GW of India's installed capacity and runs at a low plant load factor because of gas price and availability — figures change yearly, so check CEA's current monthly report.
• Emissions: gas turbine NOx and particulate limits are notified under the Environment (Protection) Act, 1986 (MoEFCC's December 2015 thermal power plant notification and its amendments) — NOx limits for gas turbines are of the order of 50–100 mg/Nm³ depending on unit size and commissioning date; verify the current amendment before quoting.
• Aviation: design, manufacture and maintenance are governed by DGCA Civil Aviation Requirements (CAR 21 for design/production organisations, CAR 145 for maintenance organisations, CAR 66 for licensed engineers), with type certification normally to FAA Part 33 / EASA CS-E standards; military engines follow CEMILAC certification.
• Indian manufacture and R&D: BHEL (Hyderabad/Bengaluru) for industrial gas turbines and HRSG-based plants, HAL Engine Division Bengaluru (licence-built Adour, AL-31FP, HTFE-25 and HTSE-1200 development), GTRE Bengaluru (Kaveri and the dry derivative), the HAL–Safran joint venture for helicopter engines, ISRO's LPSC for liquid and cryogenic rocket engines, and private suppliers in the global aero-engine supply chain.

Key points

• Gas turbine - continuous-flow prime mover: compressor, combustor, turbine; follows the constant-pressure Brayton cycle.
• Open vs closed cycle - open changes its working fluid and burns fuel internally; closed heats a fixed charge externally and is bulky but fuel-flexible.
• Work ratio - net work ÷ turbine work, only 0.3–0.5, which is why component efficiencies matter so much.
• Pressure ratio - efficiency rises with rp; net work peaks at rp,opt = (T3/T1)^γ/2(γ−1) and falls after it.
• Turbine inlet temperature - the main lever on both output and efficiency; limited by blade material, creep and cooling air.
• Regeneration - exhaust preheats compressed air; useful only at low pressure ratios; raises η, not net work.
• Intercooling - cuts compressor work, raises net work, lowers η unless regenerated.
• Reheat - raises turbine work and net work, lowers η, raises exhaust temperature.
• Combined cycle - gas turbine + HRSG + steam turbine, 52–60% net efficiency, the benchmark for gas-fired power.
• Air–fuel ratio 50:1 to 100:1 - most of the air is dilution air to keep the turbine inlet temperature safe.
• Combustor types - can, annular (lightest, aero standard), can-annular; primary, secondary and dilution zones.
• Blade cooling - convection, impingement, film and transpiration, plus thermal barrier coatings on nickel superalloys.
• Thrust - rate of change of momentum plus pressure thrust; maximum at zero speed, falls with altitude as density falls.
• Propulsive efficiency - 2/(1 + Cj/Ca); accelerate a large mass slightly, not a small mass greatly.
• Bypass ratio - high BPR turbofans give low TSFC and low noise at Mach 0.8; low BPR plus afterburner suits combat aircraft.
• Speed ranges - turboprop below ~700 km/h, turbofan Mach 0.75–0.9, turbojet Mach 1–2+, ramjet Mach 2–4, scramjet above Mach 5, rocket anywhere including vacuum.
• Specific impulse - thrust per unit weight flow of propellant (seconds); the standard figure of merit for rockets.
• ISO rating - 15 °C, 1.013 bar, 60% RH; Indian site output is derated for ambient temperature and altitude.

Formulas and examples

Example 1 — actual simple-cycle gas turbine. Air at 1 bar, 300 K; rp = 8; TIT = 1,200 K; ηc = 0.85, ηt = 0.88; cp = 1.005 kJ/kg·K, γ = 1.4; ṁ = 60 kg/s.

• t = 8^0.2857 = 1.811 → T2s = 300 × 1.811 = 543.4 K.
• Actual T2 = 300 + (543.4 − 300)/0.85 = 586.4 K; w_c = 1.005 × 286.4 = 287.8 kJ/kg.
• T4s = 1200/1.811 = 662.5 K; ideal drop 537.5 K → actual drop = 0.88 × 537.5 = 473.0 K, T4 = 727.0 K; w_t = 1.005 × 473.0 = 475.4 kJ/kg.
• w_net = 475.4 − 287.8 = 187.6 kJ/kg; q_in = 1.005 × (1200 − 586.4) = 616.7 kJ/kg.
• η = 187.6/616.7 = 30.4% (against 44.8% for the ideal cycle at the same rp); work ratio = 187.6/475.4 = 0.395; power = 60 × 187.6 = 11.26 MW.

Example 2 — optimum pressure ratio. Same T1 = 300 K, T3 = 1,200 K, ideal cycle.

• rp,opt = (1200/300)^1.4/0.8 = 4^1.75 = 11.3; then t = √4 = 2.
• w_net,max = 1.005 × 300 × (2 − 1)² = 301.5 kJ/kg; η at this point = 1 − 1/2 = 50%.
• Note that efficiency keeps rising beyond rp = 11.3 while net work falls — maximum work and maximum efficiency do not coincide.

Example 3 — adding a regenerator to Example 1. ε = 0.75, T2 = 586.4 K, T4 = 727.0 K.

• Air leaves the regenerator at T2′ = 586.4 + 0.75 × (727.0 − 586.4) = 691.9 K.
• New q_in = 1.005 × (1200 − 691.9) = 510.6 kJ/kg; w_net unchanged at 187.6 kJ/kg.
• η = 187.6/510.6 = 36.7% — a gain of 6.3 percentage points for no extra work output, and fuel saving of 106 kJ/kg.

Example 4 — turbojet thrust and efficiencies. Aircraft at 250 m/s; ṁa = 50 kg/s; ṁf = 0.9 kg/s (f = 0.018); Cj = 800 m/s; nozzle fully expanded; CV = 43,000 kJ/kg.

• F = 50 × (1.018 × 800) − 50 × 250 = 40,720 − 12,500 = 28,220 N (28.2 kN). Neglecting fuel mass: 50 × (800 − 250) = 27,500 N — about 2.5% low.
• Specific thrust = 28,220/50 = 564 N·s/kg; TSFC = 0.9/28,220 = 3.19 × 10⁻⁵ kg/N·s = 0.115 kg/N·h.
• Thrust power = 28,220 × 250 = 7.06 MW.
• ηp = 2/(1 + 800/250) = 2/4.2 = 47.6%.
• Rate of KE added = ½ × 50 × (800² − 250²) = 14.44 MW; fuel power = 0.9 × 43,000 = 38.7 MW → ηth = 37.3%.
• ηo = 0.476 × 0.373 = 17.8% (check: 6.875 MW ÷ 38.7 MW = 17.8%, using the thrust power without the fuel-mass term).

Example 5 — combined cycle and a rocket figure. A 160 MW gas turbine at 36% efficiency exhausts to an HRSG whose steam turbine gives 80 MW.

• Fuel input = 160/0.36 = 444.4 MW; total output 240 MW → plant efficiency = 240/444.4 = 54.0%.
• Steam cycle efficiency on rejected heat = 80/(444.4 − 160) = 28.1%; the formula gives 0.36 + 0.281 − 0.36 × 0.281 = 54.0% ✓.
• Heat rate = 3,600/0.54 = 6,667 kJ/kWh against about 10,000 kJ/kWh for the simple-cycle machine alone.
• Rocket: a liquid engine with ce = 2,800 m/s and ṁp = 300 kg/s gives F = 300 × 2,800 = 840 kN and Isp = 2,800/9.81 = 285 s.

Where it is used

• Power generation: combined-cycle and open-cycle stations (NTPC Dadri, Anta, Auraiya, Kawas, Faridabad; GSEG, Torrent, GIPCL in Gujarat), used for peaking and grid balancing; BHEL and Siemens Energy supply and service the machines.
• Oil, gas and process industry: gas turbine mechanical drives for pipeline and gas-lift compressors at GAIL and ONGC, refinery and fertiliser cogeneration at IOCL, RIL and IFFCO.
• Aviation: airline and air force engine maintenance and overhaul — Air India Engineering Services (AIESL), IndiGo's MRO, HAL Engine and Overhaul Divisions, with design and analysis work at the GE Aerospace, Safran, Pratt & Whitney, Rolls-Royce and Honeywell centres in Bengaluru, Hyderabad and Pune.
• Defence and space: GTRE's Kaveri and dry-Kaveri programmes, BrahMos ramjet propulsion, DRDL/RCI missile work, ISRO/LPSC liquid and cryogenic rocket engines; marine gas turbines (LM2500) on Indian Navy frigates and destroyers.
• Job roles: gas turbine performance engineer, combustion/aerothermal engineer, blade and disc stress/creep analyst, O&M and commissioning engineer at CCGT stations, engine test-cell engineer, MRO licensed engineer (DGCA CAR 66), CFD and turbomachinery design engineer.
• Auxiliary and emerging: aeroderivative gensets for data centres and process plants, microturbines, hydrogen-fired and hydrogen-blended turbine development, and supercritical CO₂ closed-cycle research.

Interview questions

• Q: Why is the work ratio of a gas turbine so low, and why does it matter? A: The compressor consumes 50–70% of the turbine work because it compresses a gas, not a liquid; so a few percent fall in compressor or turbine efficiency wipes out a large share of net output.
• Q: Does efficiency keep rising with pressure ratio? A: Ideal efficiency rises monotonically with pressure ratio, but net work peaks at rp,opt = (T3/T1)^γ/2(γ−1); in a real cycle with component losses the efficiency itself also peaks and then falls.
• Q: Why does regeneration only help at low pressure ratios? A: Heat can flow only if the turbine exhaust is hotter than the compressor delivery; at high pressure ratios compressor delivery temperature exceeds exhaust temperature and the regenerator would heat the exhaust instead.
• Q: Intercooling increases net work but can reduce efficiency — why? A: Cooler air entering the combustor needs more fuel to reach the same turbine inlet temperature, so the extra heat input can outweigh the extra work unless a regenerator recovers it.
• Q: Why is the air–fuel ratio in a gas turbine 50:1 or more when stoichiometric is about 15:1? A: Only the primary zone burns near stoichiometric; the rest of the air dilutes the gas to a temperature the turbine blades can survive and cools the liner.
• Q: What is compressor surge and how is it avoided? A: A violent flow reversal when the compressor is pushed beyond its stable operating line; avoided with variable inlet guide vanes and stators, bleed valves, multiple spools and a fuel schedule that keeps the operating point clear of the surge line.
• Q: Why do high-bypass turbofans burn less fuel than turbojets? A: They give the same thrust by accelerating a much larger mass of air to a lower jet velocity, which raises propulsive efficiency (2/(1 + Cj/Ca)) and also cuts jet noise.
• Q: A turbojet's thrust is maximum at take-off but its propulsive efficiency is zero — explain. A: Thrust is maximum at zero flight speed, but useful work is thrust × flight speed, which is zero when the aircraft is not moving, so no propulsive work is being done.
• Q: Why can a ramjet not be used for take-off? A: It relies on ram compression from forward motion, so at rest there is no pressure rise, no thrust — it must be boosted to roughly Mach 0.5–1 first.
• Q: What is specific impulse and why is it used for rockets rather than TSFC? A: Isp = thrust per unit weight flow of propellant, in seconds; a rocket carries its oxidiser, so performance is judged per unit of total propellant rather than per unit of fuel.
• Q: Why does a gas turbine lose output on a hot Indian afternoon? A: Air density falls, so mass flow falls for the same volume flow and compressor work per unit mass rises — output drops roughly 0.5–0.9% per °C above the 15 °C ISO rating.
• Q: How is a combined-cycle efficiency of about 55% possible when neither cycle alone exceeds 42%? A: The steam cycle generates extra power from the gas turbine's exhaust heat that would otherwise be rejected, so η_cc = ηgt + ηst − ηgt·ηst.

Common mistakes

• Using the ideal air-standard efficiency when component efficiencies are given — always compute actual T2 and T4 first.
• Applying the compressor's pressure-ratio temperature relation with the wrong γ or cp for the hot gas side (air γ = 1.4, cp = 1.005; gases γ ≈ 1.33, cp ≈ 1.15).
• Forgetting that the fuel mass is added to the turbine flow, or forgetting to include pressure thrust when the nozzle is choked.
• Writing thrust as ṁ(Cj − Ca) in a question that gives nozzle exit area and pressures, and losing the pressure thrust term.
• Confusing propulsive, thermal and overall efficiency, or reporting overall efficiency as their sum rather than their product.
• Assuming regeneration, intercooling and reheat all raise efficiency — alone, only regeneration does; the other two raise net work.
• Quoting thrust at sea level static as if it held at cruise altitude, and using temperatures in °C instead of kelvin in ratio formulas.

Related topics

Power cycles: Rankine, Otto, Diesel and Brayton · Turbomachinery: turbines, pumps and compressors · Air compressors · Steam turbines and nozzles · Boilers and steam power plants · Second law, entropy and availability