01 PropulsionLab · v2.6

Gas-turbine cycle analysis,
station by station.

A browser tool for turbojet, turbofan, turboprop, ramjet and scramjet cycles. Set the design point, read every station, sweep a parameter, compare engines, and export the results. The physics is reduced-order and every assumption is shown.

Engine families
5
Stations tracked
9
Solver parameters
22
Mach envelope
0→8

02 Architecture Turbojet

Select an engine architecture.

Switch architecture to load the matching educational solver. Validated turbojet includes parameter sweep and dry/reheat compare.

CONCEPT · ENGINE FAMILIES
All five families breathe air and burn fuel, they differ in how they compress, how they expand, and whether they extract shaft work. A turbojet has only a turbine to drive its compressor; a turbofan adds a fan and a bypass stream; a turboprop extracts shaft power for a propeller; a ramjet compresses with the flight Mach alone; a scramjet burns at supersonic combustor flow.
Read manual →

03 Cross-section

Hover any section to inspect the local state.

Hover a station or section to inspect the local cycle state.

04 Console

Configure the cycle and inspect results.

Parameters

Flight conditions
Engine core
Combustor
Nozzle
Afterburner

Results

Default case
Net thrust,
TSFC,
Fuel-air ratio,
Nozzle,
Variant,
Overall η,

CYCLE INSIGHTS · UPDATED EACH RUN

  • Run a cycle to populate plain-English interpretations of the numbers above.

COMBUSTOR EMISSIONS · REACTOR NETWORK

EI NOx [g/kg],
EI CO [g/kg],
Flame T [K],
φ overall,

Run a cycle to compute Cantera reactor-network emission indices.

Stn Name Tt [K] Pt [kPa] Static T [K] Static P [kPa] Mach V [m/s] Notes

Parameter sweepi

A sweep re-runs the cycle at the same design point but for a list of values of one input — everything else stays fixed. That isolates the effect of that single knob. Use it to answer questions like “does another two points of compressor PR still buy me thrust at this Tt₄?”, or “where does TSFC stop falling as I raise turbine inlet temperature?”. Sensitivity (in the Analyze tab) does this for many inputs at once and ranks them; a sweep gives you the shape of one curve.

Assumptions

  • Steady, one-dimensional gas turbine cycle.
  • Perfect gas with constant cp / γ in each gas region; separate properties for air and combustion gas.
  • Educational, preliminary-performance analysis. Not certified design software.
  • Compressor/turbine maps are synthetic illustrative characteristics sized to the design point, not measured manufacturer data.
  • Transient spool dynamics is the bare rotor-inertia response (turbojet single-spool, turbofan two-spool); no fuel-control acceleration schedule.
  • Real-gas chemistry is an optional Cantera correction across the whole cycle (variable-cp compressor, turbine and nozzle), reported beside the constant-cp deck. Default and core table stay constant-cp. No blade-row CFD.
  • Inlet capture-area mass flow is a first-order ρ·V·A estimate; no distortion or boundary-layer model.

Performance report

Builds a branded DAS LABS performance brief with the current input deck, full station table, choking state, efficiencies, and warnings. “Export Python” downloads a standalone, dependency-free script that reproduces the current run against the API.

Station stagnation temperature

Station stagnation pressure

T-s diagram

Stagnation-property cycle diagram. Entropy is on a relative datum (cp steps from air to combustion gas across the burner), so the cycle shape is meaningful but the absolute entropy zero is qualitative.

P-v diagram

Stagnation P–v (v = R·Tt/Pt), consistent across all stations, not a static-state diagram.

Sweep · thrust

Sweep · TSFC

Efficiency estimates

Thrust breakdown

Engine compare

Side-by-side run of multiple engine specs at their own design points using /compare/engines. Default deck compares the active turbojet input against representative turbofan, turboprop, ramjet, and scramjet baselines. Use the controls below to add or remove rows.

READING THE COMPARISON · THRUST ↔ TSFC

Every airbreathing engine trades the thrust it produces against the fuel it burns to produce it. TSFC (thrust-specific fuel consumption, kg/kN·hr) is just fuel flow divided by thrust, so a lower TSFC means more thrust per kg of fuel. The two numbers move in opposite directions for a deep physical reason: jet propulsion is most efficient when the engine takes a large mass of air and accelerates it a little. The same impulse produced by accelerating a small mass a lot wastes energy in the high-velocity jet that the aircraft leaves behind it. This is why the engine family ladder — turbojet → low-bypass turbofan → high-bypass turbofan → turboprop — trades top speed for fuel burn, and why ramjets and scramjets sit off that ladder entirely.

Turbojet — high thrust, high TSFC

All of the air goes through the core and leaves as a fast, hot jet. Specific thrust (N per kg/s of air) is very high, so the engine is compact and the thrust-to-weight ratio is excellent — exactly what you want at Mach 2. But the exhaust velocity is much higher than the flight velocity, so propulsive efficiency ηp = 2/(1 + Vjet/V0) is poor, and TSFC is the worst of the gas-turbine family.

Real example. The J58 on the SR-71 cruised at Mach 3 with TSFC near 1.9 kg/kN·hr — perfectly acceptable because the airframe needed thrust at that speed at all, and modern fans cannot run there. A pure turbojet earns its keep on military fighters, interceptors and supersonic transports.

Turbofan — the sweet spot for subsonic flight

A fan in front of the core accelerates a large bypass stream to a moderate velocity. Total air mass flow goes up, average jet velocity comes down, and ηp rises sharply — so TSFC drops, even though specific thrust (per kg/s of air) is lower than a turbojet's. The bypass ratio (BPR) is the dial: higher BPR trades specific thrust for lower TSFC.

Real examples. The CFM56 (BPR ≈ 5–6) on a 737 / A320 sits near TSFC ≈ 0.55 kg/kN·hr. The Pratt & Whitney GTF (BPR ≈ 12) and the Trent XWB (BPR ≈ 9) push that down to ≈ 0.49. Every BPR step costs nacelle drag and core size; the airframe sets the ceiling. Try it in PropulsionLab: open the turbofan tab and sweep bypass ratio 1 → 12 — watch thrust per kg/s of air fall and TSFC fall with it.

Turboprop — extreme bypass, extreme efficiency

Take the bypass logic to its limit: drive a large-diameter propeller, which moves an enormous mass of air at a modest velocity through the propeller disk. TSFC (when expressed against equivalent shaft power) is the best of any airbreathing engine, but the propeller's tip Mach number caps useful airspeed near M ≈ 0.6–0.7 — beyond that, shock losses on the blade tips erase the efficiency gain.

Real examples. The PW127 on an ATR 72 or the AE2100 on a C-130 burn dramatically less fuel per nautical mile than a same-size jet at their operating Mach. Try it: at the same altitude/airspeed, compare the turboprop and turbofan baselines — thrust will be comparable, but the turboprop's fuel flow is lower.

Ramjet — no compressor, no thrust until you are fast

Above roughly Mach 2.5–3, the inlet's own ram compression is enough to support combustion without a mechanical compressor. The engine is mechanically simple and survives high stagnation temperatures, but it produces no static thrust: at Mach 0 the inlet does no work, so net thrust is zero or negative. TSFC sits between turbojet and rocket — better than a rocket because oxidiser is free, worse than a turbojet because the inlet pressure recovery drops rapidly with Mach.

Real example. The D-21 drone and the SR-71's J58 in its "ramjet mode" run on this cycle above Mach 3. Try it: in the ramjet tab, sweep Mach 1 → 4 — thrust starts negative and rises sharply, but TSFC is high until ram compression is meaningful.

Scramjet — supersonic combustion, hypersonic only

Above ~Mach 5 the inlet cannot decelerate flow to subsonic without enormous total-pressure loss, so combustion has to happen in a supersonic stream. The flame-residence time is milliseconds, and dissociation chemistry above ~2500 K steals a chunk of the released heat — both of which cap practical performance. Thrust per unit air is low, TSFC is poor relative to a turbofan, but the engine is the only airbreathing option at Mach 6+.

Real example. The X-43A demonstrated scramjet flight at Mach 9.6 in 2004 for ≈ 10 s. Try it: at Mach 8, see the model warn you about dissociation as turbine-inlet temperatures approach the chemistry limit.

One mental model. If you fix the thrust requirement, the engine that burns the least fuel is the one whose jet velocity sits closest to the flight velocity. Subsonic airliner? High-bypass turbofan or turboprop. Cruise at Mach 1.6? Low-bypass turbofan. Mach 3? Turbojet. Mach 5+? You are out of options that contain a compressor at all.

User manual

PropulsionLab is an educational, station-based 1-D cycle simulator. The physics is reduced-order: perfect-gas thermodynamics, no maps, no transient dynamics, and only the chemistry that each engine module explicitly states. Every preset is labelled as an educational approximation based on public-estimated values.

Turbojet

Working principle. Single-spool gas turbine cycle. Inlet decelerates air; compressor raises Pt; combustor adds heat at approximately constant pressure; turbine extracts the work to drive the compressor; convergent nozzle accelerates the hot gas. Net thrust comes from momentum + pressure terms.

Stations. 0 freestream · 2 inlet exit · 3 compressor exit · 4 combustor exit / turbine inlet · 5 turbine exit · 7 afterburner exit (when used) · 9 nozzle exit.

Key assumptions. Steady 1-D flow; constant cp/γ in air and combustion-gas regions; energy-balance fuel-air ratio (or Cantera equilibrium when enabled). Compressor/turbine maps are offered as a separate higher-fidelity tool but are synthetic, not measured data; no blade-row CFD.

Turbofan

Working principle. Two-spool architecture. The fan pressurises both core and bypass streams. The HP spool drives the HP compressor. The LP spool drives the fan via the LP turbine. Thrust comes from a core nozzle (post-LPT) and a bypass nozzle (post-fan), or, in mixed-flow mode, from a single nozzle after a constant-pressure mixer. Optional afterburner reheats the core (separate flow) or the mixed stream (mixed flow).

Stations. 0 / 2 / 13 fan exit / 19 bypass nozzle exit / 3 HPC exit / 4 combustor exit / 45 HPT exit / 5 LPT exit / 7 AB or mixer exit / 9 core or mixed nozzle exit.

Key assumptions. No fan / compressor / turbine maps; fixed pressure ratios. Mixer is a constant-pressure mass-weighted blend with a small total-pressure loss. No bleed, no cooling flow.

Limitations. No geared-fan gearbox loss; no variable cycle; no fan distortion model.

Turboprop

Working principle. Gas-generator core (compressor / combustor / HP turbine) plus a free power turbine that drives the propeller through a reduction gearbox. Residual nozzle produces a small jet thrust on top of the propeller thrust.

Stations. 0 / 2 / 3 / 4 / 45 gas-gen exit / 5 power turbine exit / 9 residual nozzle exit.

Propeller. Efficiency comes from a Gaussian-shaped advance-ratio curve peaked at J*, with a compressibility derate above tip Mach 0.85. Below ~5 m/s freestream the model falls back to an actuator-disk static thrust estimate.

Outputs. Shaft power (SHP), equivalent shaft power (ESHP = SHP + jet thrust × V₀), BSFC in kg/(kW·h), propeller and residual jet thrust components.

Ramjet

Working principle. No rotating machinery. Inlet converts kinetic energy into compression. Combustor heats the gas at subsonic combustor Mach. Convergent nozzle accelerates the hot gas to choking and beyond.

Stations. 0 / 2 inlet exit / 3 diffuser exit / 4 combustor exit / 9 nozzle exit.

Inlet recovery. Either user value or the MIL-E-5008B curve (whichever is smaller). Above ~M5 the curve drops below ~0.5 and the model emits a fidelity warning.

Thermal choking. The combustor uses a Rayleigh-style downstream-Mach estimate; if heat addition pushes the flow to M ≈ 1 the model emits a critical thermal-choke warning.

Scramjet (reduced-order)

Working principle. Multi-shock inlet sets the supersonic combustor inlet state. Combustor adds heat at approximately constant Mach with a Rayleigh-style total-pressure loss. Convergent nozzle expands the products. Fuel input is given as an equivalence ratio φ.

Stations. 0 / 2 / 3 supersonic combustor inlet / 4 combustor exit / 9 nozzle exit.

Honest limitations. No isolator pressure recovery curve, no finite-rate chemistry, no real-gas dissociation above ~2500 K. The model is an educational reduced-order scramjet approximation, not a CFD or design tool. Warnings flag temperatures above 3000 K and below-Mach-4 operation.

Efficiency definitions

Thermal η, useful jet power (Δjet KE + pressure power) divided by fuel chemical power.

Propulsive η, thrust × V₀ divided by useful jet power. Tends to zero at static conditions by definition.

Overall η, thrust × V₀ divided by fuel chemical power. For pure-jet engines overall ≈ thermal × propulsive. For turboprop engines the overall metric includes the propeller shaft thrust, so the product rule does not hold.

Glossary, quick definitions
Station
A numbered location along the engine axis (0 freestream → 9 nozzle exit). Each station carries a thermodynamic state.
Stagnation (total) state
Properties the flow would have if brought to rest isentropically, Tt, Pt. Useful because Tt is conserved across adiabatic devices and Pt only drops where there is loss.
Static state
Local thermodynamic state of the moving gas, T, P. Differ from stagnation by the kinetic-energy term.
Mach number
Local flow speed divided by local speed of sound. Sets compressibility behaviour.
Pressure ratio (PR)
Pt at exit over Pt at inlet of a component, compressor, fan, turbine all have one.
Bypass ratio (BPR)
Mass flow through the fan bypass divided by mass flow through the core. Higher BPR → quieter, more efficient at subsonic cruise.
Fuel-air ratio (f)
Fuel mass flow divided by air mass flow through the combustor. Typically 0.015–0.04 in a lean turbojet.
Equivalence ratio (φ)
Actual fuel-air ratio divided by stoichiometric, φ < 1 is lean, φ = 1 stoichiometric, φ > 1 rich. Used for scramjets.
TSFC
Thrust-specific fuel consumption, fuel mass flow per unit thrust. Lower is better. Reported here in kg/(kN·hr).
BSFC
Brake-specific fuel consumption, fuel mass flow per unit shaft power. Used for turboprops.
Thermal efficiency
Useful jet power divided by fuel chemical power. Captures how well the cycle converts heat into kinetic energy.
Propulsive efficiency
Thrust × V₀ divided by useful jet power. Captures how well kinetic energy is turned into propulsive work.
Overall efficiency
Thermal × propulsive (for pure-jet engines). What the aircraft actually gets per unit fuel.
Choked nozzle
Nozzle throat at Mach 1, fixed mass-flow-per-area. Happens when the nozzle pressure ratio exceeds the critical value (~1.89 for hot gas).
Pressure thrust
Force from the exit-static minus ambient-static pressure times exit area. Important when the nozzle is under- or over-expanded.
Adiabatic flame temperature
Temperature reached if combustion happened with no heat loss, the thermodynamic ceiling for combustor exit temperature.
Comparison framework

Use /compare/engines or the "Engine compare" tab to run 2-5 engine specs side by side at their own design points. The output is a table with thrust, TSFC, fuel-air ratio, and the three efficiency metrics. Comparing across engine families is informative but be careful: thrust definitions and design Mach envelopes differ.

Off-design matching

Pick an engine and fix the flight condition, then compute the matched operating line. At each throttle the solver balances the spool work and choked-flow continuity to find the self-consistent point, the engine spins up and down as a fixed-geometry machine actually would. Drag the throttle to scrub thrust and TSFC along the line.

Press “Compute envelope” to match the operating line.

,
Thrust
,
TSFC
,
Pressure ratio
,
Matched
,
Solver
,

Bypass ratio (turbofan) is held at its design value; a real bypass-ratio shift needs a fan map. Low-throttle / high-Mach corners where a nozzle unchokes are dropped from the line.

Mission profile

Build a flight as a sequence of legs, each held at a fixed altitude, Mach, and throttle for a set duration. The engine is matched off-design at every leg and fuel burn is integrated across the mission. Edit any cell, add or remove legs, and the totals update automatically.

Leg Altitude m Mach Throttle, turbine inlet T K Duration s Fuel kg

Total fuel
,
Total time
,
Legs matched
,

Compressor map

Pressure ratio against corrected mass flow, with constant corrected-speed lines, surge and choke boundaries, and the matched off-design running line. The operating point is converged on the map at each throttle, drag the throttle to move it along the line. The map is sized to the deck on the Cycle tab; change the deck and recompute to resize it.

Matching the running line…

Compressor
,
Pressure ratio
,
Corrected mass flow
,
Map efficiency
,
Surge margin
,
Thrust
,
On map
,

This characteristic is a synthetic, illustrative map, correct trends, not measured manufacturer data. A verified dataset can be loaded into the same viewer later. Corrected mass flow is referenced to 288.15 K / 101.325 kPa at the compressor face; efficiency and surge margin are read off the map at the matched point.

Multi-objective design optimization · NSGA-II

Turbojet · Turbofan

A genetic algorithm (NSGA-II, non-dominated sorting, crowding distance, simulated-binary crossover and polynomial mutation, implemented from scratch) searches the design space of compressor pressure ratio and turbine-inlet temperature at the flight condition from the Cycle tab. It traces the Pareto front, the set of designs where you cannot improve one objective without giving up another, subject to a compressor-exit (material) temperature cap and a fuel-air band.

Press “Run optimization” to trace the Pareto front.

Front size
,
Evaluations
,
Min TSFC
,
Max specific thrust
,

Each point is a non-dominated design: hover the front to read it. Lower-TSFC designs push pressure ratio up until the Tt₃ cap binds; higher specific-thrust designs run the turbine hotter. The search is deterministic for a fixed seed.

Sensitivity · tornado chart

Turbojet · Turbofan

Each design input from the Cycle tab is nudged up and down by a fixed percentage; the cycle is re-run and the change in the chosen output is plotted. The longest bars are the inputs the result cares about most. This is a local, one-at-a-time analysis, it shows each input's slope and sign near the current operating point, not how inputs interact.

Press “Run sensitivity” to rank the inputs.

Baseline
,
Top driver
,

Bars extend from the baseline: blue is the change from raising the input, amber from lowering it. A long two-sided bar means a strong, monotonic driver; a one-sided or stubby bar means the output barely moves. Re-run after changing the Cycle deck.

Transient spool dynamics

Turbojet · Turbofan

Slam the throttle and watch the engine catch up. The fuel, and turbine temperature, respond almost at once, but the spinning spool has inertia, so its speed, and the thrust, lag behind. This integrates the rotor equation of motion (I·Ω·dΩ/dt = turbine power − compressor power) through the step.

Press “Run transient” to slam the throttle.

Spool τ₀
,
Settling (95%)
,
Idle thrust
,
Final thrust
,

Blue is spool speed (% of design), amber is thrust; the dashed step is the throttle command. The gap between the command and the spool is the lag. Add inertia and it spools up slower. This is the bare inertial response, a real fuel controller ramps the throttle to protect surge margin, slowing it further.

Variable geometry · reheat nozzle & flame stability

Turbojet only

Lighting the afterburner makes the exhaust much hotter, so the nozzle throat must open to pass it without backing pressure onto the turbine, and the reheat flame only stays lit inside a stability loop. Uses the turbojet deck from the Cycle tab.

Press “Analyse reheat” to schedule the nozzle.

Nozzle opens
,
Thrust augmentation
,
Dry throat area
,
Reheat throat area
,
AB equivalence φ
,
Flame state
,

The chart shows the lean-blowout equivalence ratio climbing with altitude as the afterburner-inlet pressure falls, the stability loop shrinks, and below the relight floor the burner cannot stay lit. Throat areas and stability limits are reduced-order educational estimates, not combustor CFD.

Real-engine case studies

Ten production gas-turbine engines, from the 1960s low-bypass JT8D to today's geared and three-spool high-bypass fans. Each entry gives a technical overview of the engine's architecture and cycle alongside its key specifications. All figures are public-estimated and approximate; exact numbers vary by sub-model and rating. Browse all long-form case studies →

05 Creator

Built by Ainesh Das.

Ainesh Das

Mechanical engineering undergraduate · propulsion & simulation

I'm a second-year mechanical engineering student. I built PropulsionLab to actually understand how gas-turbine cycles work, by making every station and every assumption visible instead of hiding them behind a black box. Each number comes from a real calculation, and where the model is simplified I say so. The backend is FastAPI, Pydantic and NumPy/SciPy (with optional Cantera for real-gas chemistry); the visuals are hand-written vanilla JS and SVG, no framework.

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