ThrustOne User Guide

ThrustOne supports Apple silicon Macs running macOS 12 or later, and iPads running iPadOS 16 or later.

This guide is generated from the capability catalog, field catalog, model descriptions, and demonstration cases. Demonstration results help identify controls and workflows; they do not represent a specific engine design, performance certification, or full-envelope prediction.

ThrustOne Engine Performance Calculator

Getting started Choose an engine and analysis task, adjust the conditions, then select Calculate (Command-Return). Return or leaving a field commits its value; Escape restores it. Decimal numbers, scientific notation, and the Unicode minus sign are accepted. Units appear beside fields. Internal calculations and exports use SI values. A curve and its results always belong to one complete input state. A changed input makes earlier results stale; recalculate before exporting.

Requirements ThrustOne supports Apple silicon Macs running macOS 12 or later, and iPads running iPadOS 16 or later.

Opening and saving cases Native ThrustOne documents use the .thrustone extension and contain ordinary JSON. On Mac, Open case opens a separate .thrustone or JSON document; Save case exports the current case as JSON. On iPad, use the document browser to open or create separate .thrustone cases. The Case actions menu offers Save Case As and Import into Current Case; importing accepts .thrustone or JSON, validates it first, and explicitly asks before replacing and saving the current document. Cancel keeps its existing contents. Results and chart exports use the native file picker on both platforms.

Available calculations • Turbojet: ideal and loss-based design points, afterburning, explicit-reference double-choked and dual-spool afterburning off-design points, and five standalone components. Compressor map files use shape-preserving interpolation and are never extrapolated. • Turbofan: separate-flow single-spool design, reference-normalized dual-spool separate-flow off-design, and constant-cp dual-spool mixed-flow or mixed-afterburning analyses. Constant-cp mixing matches inlet static pressure and uses the subsonic conservation root. NASA9 variable-cp design analyses include separate flow, mixed flow, and independent post-mixer afterburning. Specific thrust uses total inlet air; fuel-air ratio uses core air. Ideal modes neglect fuel mass addition and assume fully expanded streams. • Liquid rocket: thrust-chamber and nozzle performance with explicit gas properties; altitude, area-ratio, and chamber-pressure sweeps at fixed reference scale; equivalent cylindrical sizing; cylinder-plus-convergent volume allocation; a standalone finite-contraction static chamber plane; and conical or bell reference geometry. Mixture-ratio sweeps never combine values based on different chemical assumptions. • Ramjet: full-engine design point, normal shock, explicit weak or strong oblique shock, Darcy Fanno friction, same-branch Rayleigh heating, and fixed-cp quasi-one-dimensional flow. Flow segments support area, Darcy friction, heat flow, and injected mass. They stop at sonic events or residual limits and retain valid history. • Solid rocket: equilibrium pressure from a supplied burn-rate law and end-burning or internal-port regression transients. Demonstration parameters do not represent a specific propellant. Models start from an established choked state and exclude ignition, erosion, and tail-off blowdown. • TBCC: independent turbojet and ramjet channels combined with an explicit air split. The shared freestream is defined in turbojet fields. Common-inlet matching, variable geometry, and automatic transition are not solved. • RBCC: an ideal ejector-ramjet submodel with specified primary total state, supersonic primary flow, subsonic secondary flow, and a choked ejector exit. It is not a full multimode RBCC cycle. • ATR: an ideal air-turboramjet submodel with specified primary total state, ideal work balance, low-speed mixing, and simplified heating. It excludes a gas generator, pump, and cooling.

Freestream and model limits Turbojet input supports Mach plus ambient state, altitude plus Mach, Mach plus dynamic pressure, and altitude plus dynamic pressure. Altitude is geometric; the standard atmosphere is valid from 0 to 86 km. Legacy JSON altitude values in km convert to m on import. The dynamic-pressure path uses standard air with γ = 1.4.

Ideal cycles use one gas and neglect fuel mass addition. Other constant-cp cycles use supplied cold- and hot-side heat capacities and component losses. Constant-cp calculation is not chemical equilibrium. Liquid-rocket empirical efficiencies correct characteristic velocity or thrust coefficient; they do not change the ideal reference exit velocity. Exit-plane static pressure is not a validated wall-pressure or side-load prediction.

Cases and files Open and save native JSON cases up to 10 MiB and 64 nesting levels. Version 2 uses engine, mode, and the associated input section. Unsupported modes report an error; they are never silently replaced. Some legacy turbojet cases and rocket cases with explicit reference conditions can import; other thermodynamic-table branches have not been migrated.

Edits in JSON Interchange update the form only after Apply JSON. Restore Form Contents discards unapplied text, which is not written to the saved case. Unknown fields are retained as extension data; known fields in this version determine the calculation. The File menu provides Save, Save As, Close, and multiple windows.

Studies and exports Parameter Study operates on currently exposed independent fields. It accepts linear endpoints or explicit value lists, one- or two-dimensional grids, and at most 1,000 total cases. It shows progress and can be cancelled while retaining completed work. Failed rows retain their error. Liquid-rocket plots do not join across failed points. History tables show the first 100 points; JSON, CSV, and graphic exports retain all data.

Study axes, explicit lists, and the two-dimensional switch are saved with a case. The comparison metric and maximum/minimum choice are view state. Grid extrema cover only successful computed points; they are neither continuous nor global optima. A three-point curve may show an endpoint secant slope; reduce the step to check local stability.

Metric Constraints and Comparison Report apply inclusive SI lower and upper bounds to existing results; every selected constraint must pass. Apply the constraints first. Reapplying the same metric replaces its bound; Remove removes it. Rows with missing metrics or errors cannot qualify. The report records the request, axes, constraints, extrema before and after filtering, and each row’s status. Filtering does not rerun the solver and is not continuous optimization or a complete performance envelope.

One-dimensional studies provide local golden-section search. Inspect objective shape with a grid first, specify an interval, and declare that it can be treated as unimodal there. The search uses 4–100 evaluations and stops on combined absolute and relative interval-width tolerances. An out-of-domain, non-finite, missing-metric, or constraint-violating point stops it; that point is not skipped. The result is only the best evaluated point. Budget exhaustion is not convergence, and interval convergence neither proves a global optimum nor bounds objective error. Settings are view state; full conditions and evaluated points export as a separate JSON record.

A history plot can overlay one same-unit series: blue circles are the primary quantity and orange squares the overlay. The legend and numeric table show both. Each series retains its gaps; failed points are not joined. Different dimensions never share a forced vertical axis. A two-dimensional grid shows a discrete feasible set: passing points are colored by objective, red crosses retain solver failures, and gray diamonds retain rejected or missing-metric points. It does not interpolate or draw a continuous envelope. PNG and PDF export are available; the comparison report contains all raw rows and reasons.

Result JSON contains the case used, model mode, SI values, and histories. CSV uses UTF-8 with BOM and CRLF, with every quantity’s unit and an input snapshot. PNG is 1000 × 600; PDF is vector graphics. Stale results cannot be exported as current results.

Offline and privacy Calculations run locally. ThrustOne does not connect to remote services, upload cases, or use model APIs. Files are saved only where you choose.

Scope ThrustOne provides 49 runnable calculation models across turbojets, turbofans, liquid and solid rockets, ramjets, scramjets, TBCC, RBCC, and ATR. The capability catalog and current mode description state each model’s conditions and limits. Interpret results only within the selected model. Demonstration inputs explain operation; they do not represent a specific engine design, performance certification, or full-envelope prediction.

The form footer can switch display between SI and engineering units (kPa, kN, kJ); primary results convert and the preference is saved with the case. Commit any field being edited before switching. The calculation core, JSON, CSV, station data, curves, and explicit lists always use the labeled SI values. Temperature remains K.

Thermochemistry and variable-property models Scramjet full-engine, combustor, and fixed-cp frozen-nozzle modes use explicit heat and mass sources and stop at sonic events. In seven-state fixed-cp F/G input, F = pA + ṁV in N and G = cpT + V²/2 in J/kg. Do not enter reacting-flow total enthalpy containing species formation enthalpies in that form.

NASA9 properties use archived U.S. government species data (NASA/TP-2002-211556) only in the common 200–6000 K range of selected species. Enthalpy includes formation enthalpy; pressure is Pa and molar mass is kg/kmol. Gas equilibrium uses 13 species with no ions or condensed phases. Displayed cp and γ are local frozen-composition properties.

Variable-cp frozen nozzles solve enthalpy, entropy, and mass continuity. Thermochemical rocket tasks can solve chamber equilibrium at specified chamber temperature and pressure, then expand with frozen composition. Chamber temperature can also close from gaseous inlet enthalpy and equilibrium combustion. Along-path equilibrium and user-selected downstream-area freezing are available. These models exclude liquid-propellant vaporization and condensed phases.

Finite-rate reactor mode is constant-volume and externally isothermal. You supply reactions and Arrhenius coefficients. A uses kmol/m³ concentrations and K; Ea uses J/kmol. Default coefficients demonstrate the algorithm and are not a combustion mechanism. Reverse rates derive from selected thermodynamic data. The supplied hydrogen–oxygen 10-species/29-reaction option includes third-body and Lindemann/Troe corrections. Stiffness or step limits retain an incomplete state, which cannot participate in study extrema.

The ramjet path’s final nozzle uses specified exit pressure and back-calculates area. It shows two stations and is not fixed-geometry back-pressure matching. A bell nozzle is arc–parabola geometry with explicit endpoint angles; it is not a validated wall-pressure, separation, or side-load prediction.

Constant-pressure energy closure uses gaseous inlet composition, common inlet temperature, and heat per unit mixture mass to find equilibrium temperature. Zero heat is adiabatic. Every temperature trial must satisfy elemental conservation and equilibrium residuals; an unbracketed user temperature interval is rejected. There is no liquid vaporization, condensed phase, or ion model.

Thermochemical rocket analysis permits specified chamber temperature or inlet-energy closure. Gaseous mixture-ratio sweeps use H2/CH4/CO with O2; MR is oxidizer-to-fuel mass ratio. Every point keeps geometry and back pressure fixed and uses the same chamber-equilibrium and selected nozzle-chemistry model. Best Isp means the maximum among successful samples; failed rows remain and no continuous optimum is inferred.

Variable-cp turbojet and separate-flow turbofan components use NASA9 frozen composition and h(T), s(T,p) inversion. Inlet uses supplied total-pressure recovery only. Compressor and turbine use separately entered isentropic efficiencies; do not treat a previous polytropic efficiency as the same parameter. Nozzles expand to specified exit static pressure and do not perform fixed-geometry matching. This path has no compressor map or combustion reaction. Fan, booster, and bypass nozzle share these components; the mixer has its selected mixing model.

NASA9 turbojet and separate-flow turbofan design points use dry air at 79 mol% N2 and 21 mol% O2. Gaseous H2/CH4/CO fuel is specified by fuel-to-core-air mass ratio, and fuel inlet temperature is included separately in formation enthalpy. The combustor solves adiabatic gas-phase equilibrium, then freezes composition. The turbine drives core-compressor and all-air fan loads. Net specific thrust uses total air; fuel-air ratio uses core air. These modes use fully expanded variable-area exits, not fixed geometry, maps, or off-design matching. They exclude liquid vaporization, cooling, and condensation.

The calibrated variable-property inlet supports altitude plus Mach, Mach plus dynamic pressure, and altitude plus dynamic pressure with contraction ratio, exit Mach, exit static pressure, exit static temperature, or total-pressure recovery. It uses NASA9 dry air and 0–70 km geometric altitude. A contraction inversion explicitly chooses the higher- or lower-Mach root. It does not extrapolate higher cold-atmosphere properties or promise point-for-point agreement with another mechanism database.

NASA9 frozen quasi-one-dimensional flow updates heat capacity and sound speed at every step and checks total-energy residual. Injection must have main-stream composition; injection total enthalpy includes formation enthalpy and may be physically negative. After changing composition, select Update Mass Flow from Inlet State. Sonic, property-range, and residual events retain a valid prefix, which cannot be used for study extrema. It does not mix different compositions or solve finite-rate chemistry.

NASA9 seven-state input supports T/P/M, T/P/V, T/ρ/M, T/ρ/V, P/ρ/M, P/ρ/V, and F/G/ṁ. Here F = pA + ṁV and G = h + V²/2; G includes formation enthalpy. F/G requires an explicit subsonic or supersonic root.

Finite-rate reacting flow integrates species mass flow, momentum, and total energy, so temperature changes with reaction. Each segment may specify injection composition; absent a value, inlet composition is used. Injection total enthalpy includes formation enthalpy, so external heat must not add fuel reaction heat again. Default dissociation coefficients demonstrate the algorithm only. Supply reliable mechanism parameters for a physical study. The model omits transport, diffusion, and automatic transonic matching.

The calibrated full engine connects a dry-air inlet, combustor, and nozzle in either nonreacting mixing or finite-rate mode. Net thrust subtracts captured-air flight momentum. specific_impulse_injected uses all injected mass; air or oxidizer is not automatically called fuel. Geometry is supplied by the user; ambient back-pressure closure, shock trains, and mode transition are not automatic. If any segment is incomplete, full-engine thrust is not reported.

Equilibrium nozzles follow an isentropic path with the same elemental inventory; critical mass flow is found from the internal maximum of ρV. Selected-area freezing switches to fixed composition at the user’s post-throat area ratio. The local frozen Mach there must exceed 1; equilibrium throat state is not automatically a frozen-sonic throat state. The selected point is not a calculated chemical-freeze location. frozen_mach uses local frozen sound speed, not equilibrium sound speed. Property extrapolation is rejected. Current models contain ideal gas only: no liquid vaporization, condensation, wall pressure, post-separation performance, or side load.

The optional hydrogen–oxygen submechanism has 10 species and 29 reactions, including HO2/H2O2, third-body efficiencies, and Troe reactions. Source cm/mol/s and cal/mol values are converted to m/kmol/s and J/kmol. It is neither a full GRI-Mech nor a point-for-point reproduction of other software. NASA9 and hydrogen–oxygen NASA7 are independent selections. The H2/O2 mechanism common range is 300–3500 K. A standalone flow stops outside it or at the integration budget; calibrated full-engine analysis uses NASA9.

The reaction editor can create elementary, third-body, or pressure-falloff reactions. Unlisted collision efficiencies equal 1. Pressure falloff uses Lindemann with optional three- or four-parameter Troe. Units of A vary with reaction order and temperature exponent; check displayed units after changing type. Load replaces reactions, initial composition, and reactor controls. The included data are adapted from the Cantera hydrogen–oxygen submechanism; its BSD license appears under Third-Party Data Licenses. Presets only enter the workflow and do not replace experimental or textbook validation for a specific engine.

Mixture-ratio sweeps can use chamber-exit freezing, along-path equilibrium, or selected-area freezing. Every point uses the same selected model, geometry, and back pressure; older files default to chamber-exit freezing. Failed samples retain their reason. The highest discrete specific impulse is not a continuous optimum.

Variable-cp turbojet and separate-flow turbofan models offer optional core-stream afterburning. Afterburner fuel-air ratio is per core-air mass and uses the same gaseous fuel as the main combustor. Fuel inlet temperature enters formation enthalpy separately; lower heating value is not added again. The afterburner follows the turbine and does not change shaft loads; separate-flow bypass does not afterburn. Total fuel flow and specific fuel consumption include main combustor and afterburner. This mode excludes mixed flow, liquid vaporization, cooling, and map matching.

The turbofan two-stream mixer can use NASA9 variable-property frozen mixing. Enter both streams’ total states, mass flows, compositions, and core inlet Mach. Bypass static pressure matches the core, and areas are recovered from both mass flows; a constant-area exit uses only the subsonic conservation root. Composition mixes by elemental and species mass flow without chemical reaction. Total-pressure recovery is a separate loss downstream of the mixer and does not alter its momentum conservation. This component is not a mixed-flow full-engine match or FAIR/VSH validation.

Offline theory manual Choose Open Theory Manual in Help to read the bundled English reference. Search for a term, move between matches, follow the PDF contents links, or enter a page number. No internet connection is required. Each chapter states its assumptions, equations and interpretation limits.

Current Mode Reference

Turbojet · Variable-specific-heat gaseous-fuel design point

Mode identifier: turbojet.variable_cp_design.

NASA9 dry air and gaseous fuel with adiabatic gas-phase equilibrium combustion, frozen turbine/nozzle composition, optional core-flow afterburning, and a single-spool variable-area exit.

NASA9 dry air and gaseous fuel with inlet temperatures included separately in enthalpy. Adiabatic gas-phase equilibrium sets combustor temperature, then composition freezes. Single-spool load and net thrust are closed with fully expanded, variable-area exits. Optional post-turbine core-stream afterburning uses the same balance model for fuel mass and formation enthalpy; separate-flow bypass does not afterburn. No liquid vaporization, cooling, or map matching.

(1+f) h4 = h3 + f hf；(1+f)(h4−h5)ηm = wc+(1+α)wf；(1+f+fAB)h7 = (1+f)h5+fAB hfAB；F/ṁcore = (1+f+fAB)V9+αV19−(1+α)V0

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - NASA9 variable specific heat with dry air (79 mol% N2, 21 mol% O2) and gaseous H2/CH4/CO fuel. Fuel-air ratio uses core air mass; net specific thrust uses total inlet air mass. - The adiabatic combustor solves 13-species gas-phase equilibrium. Fuel inlet temperature and formation enthalpy enter the energy balance; no separate heating value is added. Composition is frozen downstream. Excludes condensed phases, liquid vaporization, afterburning, cooling, and finite-rate chemistry. - Single-spool work balance with fully expanded streams and variable exit areas. This separate-flow model does not perform fixed-geometry, component-map, mixed-flow, or off-design matching.

Turbojet · Design point

Mode identifier: turbojet.design_point.

Constant-specific-heat cycle with specified compressor pressure ratio and turbine-inlet total temperature.

Constant-specific-heat cycle with specified compressor pressure ratio and turbine-inlet total temperature. Steady, one-dimensional turbojet cycle. The ideal mode uses one gas and neglects fuel mass addition; the loss mode uses constant specific heats in separate cold and hot sections. Off-design operation is limited to the fixed-area, double-choked model.

τr = 1 + (γc − 1) M₀² / 2
τc = πc^((γc − 1)/(γc ec))
S = f / (F/ṁ₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The outlet is not fully expanded; the current kinetic energy decomposition does not include pressure thrust work, so no thermal or propulsive efficiency is reported. The total efficiency is still calculated based on thrust power/fuel chemical power.

Turbojet · Afterburning design point

Mode identifier: turbojet.design_point_ab.

Applies afterburner heat addition and total-pressure loss downstream of the turbine.

Applies afterburner heat addition and total-pressure loss downstream of the turbine. Steady, one-dimensional turbojet cycle. The ideal mode uses one gas and neglects fuel mass addition; the loss mode uses constant specific heats in separate cold and hot sections. Off-design operation is limited to the fixed-area, double-choked model.

τr = 1 + (γc − 1) M₀² / 2
τc = πc^((γc − 1)/(γc ec))
S = f / (F/ṁ₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The step-by-step constant specific heat is used for afterburning, and the cross-section enthalpy reference may jump; it does not represent variable specific heat or chemical equilibrium results. - The outlet is not fully expanded; the current kinetic energy decomposition does not include pressure thrust work, so no thermal or propulsive efficiency is reported. The total efficiency is still calculated based on thrust power/fuel chemical power.

Turbojet · Component calculation

Mode identifier: turbojet.component.

Independent inlet, compressor, combustor, turbine, nozzle, and compressor-map calculations.

Standalone component calculation using the current component inlet state and loss parameters. It does not represent a complete engine cycle or automatic component matching.

wc = cp (Tt,out − Tt,in)
τc = πc^((γ − 1)/(γ ec)) (without a compressor map)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Turbojet · Off-design point

Mode identifier: turbojet.off_design.

Normalized fixed-area, double-choked reference model; compressor efficiency and turbine temperature and pressure ratios retain their reference values.

Normalized fixed-area, double-choked reference model; compressor efficiency and turbine temperature and pressure ratios retain their reference values. Steady, one-dimensional turbojet cycle. The ideal mode uses one gas and neglects fuel mass addition; the loss mode uses constant specific heats in separate cold and hot sections. Off-design operation is limited to the fixed-area, double-choked model.

τr = 1 + (γc − 1) M₀² / 2
τc = πc^((γc − 1)/(γc ec))
S = f / (F/ṁ₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The outlet is not fully expanded; the current kinetic energy decomposition does not include pressure thrust work, so no thermal or propulsive efficiency is reported. The total efficiency is still calculated based on thrust power/fuel chemical power.

Turbojet · Dual-spool afterburning off-design point

Mode identifier: turbojet.dual_spool_afterburning_off_design.

Normalized dual-spool reference point with fixed turbine-ratio and choking assumptions; the afterburner can be disabled.

Normalized dual-spool reference point with fixed turbine-ratio and choking assumptions; the afterburner can be disabled. Steady, one-dimensional turbojet cycle. The ideal mode uses one gas and neglects fuel mass addition; the loss mode uses constant specific heats in separate cold and hot sections. Off-design operation is limited to the fixed-area, double-choked model.

τr = 1 + (γc − 1) M₀² / 2
τc = πc^((γc − 1)/(γc ec))
S = f / (F/ṁ₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Dual-spool reference normalization holds turbine temperature and pressure ratios fixed and assumes choking at stations 4, 4.5, and 8. Variable nozzle area isolates afterburner changes from turbine matching. Compressor maps and unchoked operation are not modeled.

Turbofan · Variable-specific-heat single-spool design point (separate-flow or optional mixed-flow)

Mode identifier: turbofan.variable_cp_design.

NASA9 gaseous fuel; gas-phase equilibrium combustion and frozen chamber exit; single-spool separate-flow cycle with fully expanded streams. Optional core-stream afterburning, or explicit equal-static-pressure, constant-area, subsonic conservation mixing with independent post-mixer afterburning. An incompatible mixing root is rejected; pressure loss is not used to fabricate a match. No map matching, fixed-geometry back-pressure matching, or FAIR/VSH reproduction is claimed.

NASA9 dry air and gaseous fuel with inlet temperatures included separately in enthalpy. Adiabatic gas-phase equilibrium sets combustor temperature, then composition freezes. Single-spool load and net thrust are closed with fully expanded, variable-area exits. Optional post-turbine core-stream afterburning uses the same balance model for fuel mass and formation enthalpy; separate-flow bypass does not afterburn. No liquid vaporization, cooling, or map matching.

(1+f) h4 = h3 + f hf；(1+f)(h4−h5)ηm = wc+(1+α)wf；(1+f+fAB)h7 = (1+f)h5+fAB hfAB；F/ṁcore = (1+f+fAB)V9+αV19−(1+α)V0

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - NASA9 variable specific heat with dry air (79 mol% N2, 21 mol% O2) and gaseous H2/CH4/CO fuel. Fuel-air ratio uses core air mass; net specific thrust uses total inlet air mass. - The adiabatic combustor solves 13-species gas-phase equilibrium. Fuel inlet temperature and formation enthalpy enter the energy balance; no separate heating value is added. Composition is frozen downstream. Excludes condensed phases, liquid vaporization, afterburning, cooling, and finite-rate chemistry. - Single-spool work balance with fully expanded streams and variable exit areas. This separate-flow model does not perform fixed-geometry, component-map, mixed-flow, or off-design matching.

Turbofan · Turbofan component calculation

Mode identifier: turbofan.component.

Standalone fan, booster, bypass-nozzle, and two-stream mixer calculations with constant-cp or NASA9 frozen properties. Total states and downstream additional loss are explicit inputs.

Standalone fan, booster, bypass-nozzle, and two-stream static-pressure-matched mixer calculations. The mixer may use constant-cp or NASA9 frozen species and does not solve chemical reaction. A constant-area outlet uses the subsonic conservation root; additional pressure loss acts separately downstream. The full-engine operating point is not closed automatically.

wc = cp (Tt,out − Tt,in)
τc = πc^((γ − 1)/(γ ec)) (without a compressor map)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Independent Fan calculation using the shared compression/nozzle solver; it does not perform full-engine matching.

Turbofan · Single-spool separate-flow design point

Mode identifier: turbofan.separate_design.

Separate core and bypass streams with single-spool load balance.

Separate core and bypass streams with single-spool load balance. Separate-flow, single-spool turbofan design point, with core pressure ratio measured from the inlet exit. Total specific thrust is normalized by total inlet airflow and fuel-air ratio by core airflow. The ideal mode neglects fuel mass addition. Dual-spool separate-flow off-design and constant-specific-heat dual-spool mixed-flow submodels are also available; their assumptions are listed with their inputs and results.

F/ṁtotal = (Fcore/ṁcore + α Fbypass/ṁbypass)/(1+α)
S = ṁfuel/F

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - At least one outlet is not fully expanded; the current kinetic energy decomposition does not include pressure thrust work, so thermal efficiency or propulsion efficiency is not reported. The total efficiency is still calculated based on thrust power/fuel chemical power.

Turbofan · Dual-spool separate-flow off-design point

Mode identifier: turbofan.separate_off_design.

Independent dual-spool reference-normalized matching with separate convergent nozzles and explicit reference conditions.

Independent dual-spool reference-normalized matching with separate convergent nozzles and explicit reference conditions. Separate-flow, single-spool turbofan design point, with core pressure ratio measured from the inlet exit. Total specific thrust is normalized by total inlet airflow and fuel-air ratio by core airflow. The ideal mode neglects fuel mass addition. Dual-spool separate-flow off-design and constant-specific-heat dual-spool mixed-flow submodels are also available; their assumptions are listed with their inputs and results.

F/ṁtotal = (Fcore/ṁcore + α Fbypass/ṁbypass)/(1+α)
S = ṁfuel/F

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Reference-normalized approximation with two choked HP-turbine stations, equal fan and LP-compressor outlet conditions, and fixed-throat convergent nozzles. Excludes mixed-flow matching, component maps, and surge prediction.

Turbofan · Constant-specific-heat dual-spool mixed-flow

Mode identifier: turbofan.mixed_constant_cp.

Constant-cp dual-spool mixed-flow cycle with subsonic conservation mixing and optional afterburning; no FAIR/VSH thermochemistry.

Constant-cp dual-spool mixed-flow cycle with subsonic conservation mixing and optional afterburning; no FAIR/VSH thermochemistry. Separate-flow, single-spool turbofan design point, with core pressure ratio measured from the inlet exit. Total specific thrust is normalized by total inlet airflow and fuel-air ratio by core airflow. The ideal mode neglects fuel mass addition. Dual-spool separate-flow off-design and constant-specific-heat dual-spool mixed-flow submodels are also available; their assumptions are listed with their inputs and results.

F/ṁtotal = (Fcore/ṁcore + α Fbypass/ṁbypass)/(1+α)
S = ṁfuel/F

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Constant-specific-heat dual-spool mixed-flow approximation with matched inlet static pressures and a subsonic mixer-exit root. It does not solve variable-specific-heat thermochemistry. - The mixing pressure loss is used downstream with the total pressure recovery coefficient; ideal_mixer_exit_mach is a conserved mixing state without additional losses and does not represent the local Mach after additional pressure losses.

Liquid Rocket · Thermochemical chamber and equilibrium nozzle

Mode identifier: liquid_rocket.equilibrium_thrust_chamber.

Gas-phase equilibrium isentropic expansion with the same elemental inventory. Critical mass flux closes the throat state; composition can freeze after a selected downstream area ratio. No liquid or condensed phase.

NASA9 ideal-gas, gas-phase equilibrium isentropic expansion. Critical mass flux is found from the maximum along the path; composition may freeze at a selected location. frozen_mach uses local frozen sound speed and is not equilibrium-sound-speed Mach. No liquid phase, condensed phase, or post-separation performance.

s_eq(T,p,b)=sc; V²/2=hc−h_eq; ṁ=At max(ρV); A=ṁ/(ρV); after freezing xᵢ=constant

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Ideal-gas chemical equilibrium is solved along the chamber and nozzle with a conserved elemental inventory. - Critical mass flux is the interior maximum of ρV along the equilibrium isentrope. frozen_mach uses the local frozen speed of sound; it is not the equilibrium Mach number. - Linear variation of area along the axis is used for display, without accounting for wall pressure, post-separation properties, liquid vaporization, or condensed phase. - Chamber temperature can be specified or solved from gaseous inlet enthalpy at constant pressure; data version: NASA9-c915968268c707d69776909cadd834f1bcb6e44e4aaf1adfc6d5f42f86f6ace4.

History data: - equilibrium_nozzle_profile: 31 points; charts can select a quantity and JSON/CSV exports retain complete data.

Liquid Rocket · Thermochemical nozzle · Specified freeze area

Mode identifier: liquid_rocket.partial_frozen_thrust_chamber.

Gas-phase equilibrium isentropic expansion with the same elemental inventory. Critical mass flux closes the throat state; composition can freeze after a selected downstream area ratio. No liquid or condensed phase.

NASA9 ideal-gas, gas-phase equilibrium isentropic expansion. Critical mass flux is found from the maximum along the path; composition may freeze at a selected location. frozen_mach uses local frozen sound speed and is not equilibrium-sound-speed Mach. No liquid phase, condensed phase, or post-separation performance.

s_eq(T,p,b)=sc; V²/2=hc−h_eq; ṁ=At max(ρV); A=ṁ/(ρV); after freezing xᵢ=constant

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Gas-phase equilibrium is maintained in the chamber and up to the specified freeze area; composition is frozen downstream. The chemical freeze location is a user input, not a prediction. - Critical mass flux is the interior maximum of ρV along the equilibrium isentrope. frozen_mach uses the local frozen speed of sound; it is not the equilibrium Mach number. - Linear variation of area along the axis is used for display, without accounting for wall pressure, post-separation properties, liquid vaporization, or condensed phase. - Chamber temperature can be specified or solved from gaseous inlet enthalpy at constant pressure; data version: NASA9-c915968268c707d69776909cadd834f1bcb6e44e4aaf1adfc6d5f42f86f6ace4.

History data: - equilibrium_nozzle_profile: 61 points; charts can select a quantity and JSON/CSV exports retain complete data.

Liquid Rocket · Thermochemical chamber and frozen nozzle

Mode identifier: liquid_rocket.thermochemical_frozen_nozzle.

Chamber temperature may be specified or closed from gaseous inlet enthalpy, followed by a NASA9 frozen nozzle. No liquid vaporization.

Chamber temperature may be specified or closed from inlet gas-phase energy, then expanded with NASA9 chamber equilibrium or supplied frozen composition. The throat closes by enthalpy and sound speed. No liquid vaporization or along-path reaction.

h(T*) + a*²/2 = hc；s* = sc；ṁ = ρ* a* At；c* = pc At/ṁ；Isp = F/(ṁg₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The chamber solves 13-species gas-phase equilibrium. Nozzle expansion holds the chamber-exit composition frozen; it does not follow an equilibrium expansion path. - Chamber temperature is specified by the user; inlet energy closure is not solved. - Does not include finite rate losses, condensed phase, or post-separation properties. - NASA9 data version: NASA9-c915968268c707d69776909cadd834f1bcb6e44e4aaf1adfc6d5f42f86f6ace4.

History data: - nozzle_profile: 101 points; charts can select a quantity and JSON/CSV exports retain complete data.

Liquid Rocket · Finite contraction · Static chamber plane

Mode identifier: liquid_rocket.finite_contraction.

Standalone static chamber plane, finite-contraction subsonic root, and lossless total-state conversion; no heat-addition pressure loss.

A static chamber plane and finite contraction ratio define the subsonic state, followed by lossless total-state conversion. No heat-addition pressure loss.

Tt = T [1 + (γ − 1) M²/2]
Pt = p (Tt/T)^(γ/(γ − 1))
ṁ = ρ u Ac

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Independent static chamber plane input, the total state is obtained through lossless isentropic relationship; it is not applied to other thrust chamber modes. - Finite-velocity heat-addition losses may be significant for contraction ratios below 2; Rayleigh losses and combustion efficiency are not calculated.

Liquid Rocket · Cylindrical chamber and convergent-section sizing

Mode identifier: liquid_rocket.convergent_chamber_geometry.

L*, contraction ratio, and cone half-angle allocate cylindrical and convergent sections; evaporation and combustion efficiency are not predicted.

Characteristic length sets total chamber volume; contraction ratio and conical half-angle allocate cylindrical and convergent sections. Combustion efficiency is not predicted.

Vc = L* At
Vc = V cylinder + V convergence section

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The geometric assignment of cylinders plus right cones only does not predict droplet evaporation, combustion completeness, or injector structure.

History data: - chamber_convergent_geometry: 3 points; charts can select a quantity and JSON/CSV exports retain complete data.

Liquid Rocket · Thrust-chamber performance

Mode identifier: liquid_rocket.thrust_chamber.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale. Given the gas temperature, specific heat ratio and molar mass, quasi-one-dimensional isentropic expansion is used. Empirical efficiency only corrects the coefficient of performance; outlet static pressure cannot be used as a verified prediction of wall pressure or side load.

c* = pc At/ṁ
CF = F/(pc At)
Isp = F/(ṁ g₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Liquid Rocket · Standalone nozzle performance

Mode identifier: liquid_rocket.nozzle_performance.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale. Given the gas temperature, specific heat ratio and molar mass, quasi-one-dimensional isentropic expansion is used. Empirical efficiency only corrects the coefficient of performance; outlet static pressure cannot be used as a verified prediction of wall pressure or side load.

c* = pc At/ṁ
CF = F/(pc At)
Isp = F/(ṁ g₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Liquid Rocket · Equivalent chamber sizing

Mode identifier: liquid_rocket.chamber_sizing.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale.

Characteristic length and throat area set chamber volume and equivalent cylindrical dimensions. No convergent-section allocation.

Vc = L* At
Ac = CR At
Leq = Vc/Ac

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Equivalent cylinder length of length Vc/Ac, excluding convergence section assignment, injector structure and finite contraction total pressure correction.

Liquid Rocket · Nozzle contour · Conical or bell

Mode identifier: liquid_rocket.nozzle_contour_design.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale.

Conical-nozzle reference contour defined by throat area, exit area ratio, and half-angle.

rt = √(At/π)
re = rt √ε
L = (re − rt)/tan θ

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Geometry reference of right cone expansion section only, does not include throat arc, bell shape, wall pressure, separation position or side load. The divergence factor is reported separately and is not applied again to the prescribed performance efficiency.

History data: - conical_reference_contour: 101 points; charts can select a quantity and JSON/CSV exports retain complete data.

Liquid Rocket · Altitude performance

Mode identifier: liquid_rocket.altitude_characteristic.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale. Given the gas temperature, specific heat ratio and molar mass, quasi-one-dimensional isentropic expansion is used. Empirical efficiency only corrects the coefficient of performance; outlet static pressure cannot be used as a verified prediction of wall pressure or side load.

c* = pc At/ṁ
CF = F/(pc At)
Isp = F/(ṁ g₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The sweep keeps the baseline throat dimensions and explicit gas parameters fixed; it does not automatically reset thrust, look up tables, or change the chemical model. Failure points are retained and curves are not connected across failure points.

History data: - altitude_characteristic: 101 points; charts can select a quantity and JSON/CSV exports retain complete data.

Liquid Rocket · Area-ratio sweep

Mode identifier: liquid_rocket.area_ratio_sweep.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale. Given the gas temperature, specific heat ratio and molar mass, quasi-one-dimensional isentropic expansion is used. Empirical efficiency only corrects the coefficient of performance; outlet static pressure cannot be used as a verified prediction of wall pressure or side load.

c* = pc At/ṁ
CF = F/(pc At)
Isp = F/(ṁ g₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The sweep keeps the baseline throat dimensions and explicit gas parameters fixed; it does not automatically reset thrust, look up tables, or change the chemical model. Failure points are retained and curves are not connected across failure points.

History data: - area_ratio_sweep: 101 points; charts can select a quantity and JSON/CSV exports retain complete data.

Liquid Rocket · Chamber-pressure sweep

Mode identifier: liquid_rocket.throttle_characteristic.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale.

Explicit gas properties; geometry is an equivalent chamber or conical reference. Sweeps retain a fixed reference scale. Given the gas temperature, specific heat ratio and molar mass, quasi-one-dimensional isentropic expansion is used. Empirical efficiency only corrects the coefficient of performance; outlet static pressure cannot be used as a verified prediction of wall pressure or side load.

c* = pc At/ṁ
CF = F/(pc At)
Isp = F/(ṁ g₀)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The sweep keeps the baseline throat dimensions and explicit gas parameters fixed; it does not automatically reset thrust, look up tables, or change the chemical model. Failure points are retained and curves are not connected across failure points.

History data: - throttle_characteristic: 101 points; charts can select a quantity and JSON/CSV exports retain complete data.

Liquid Rocket · Gaseous-propellant mixture-ratio sweep

Mode identifier: liquid_rocket.mixture_ratio_sweep.

Gaseous H2/CH4/CO–O2 mixture-ratio sweep using one NASA9 chamber-equilibrium model and the explicitly selected frozen, equilibrium, or selected-area-freezing nozzle model. Chamber temperature is fixed or set by gas-phase energy closure.

Gaseous H2/CH4/CO–O2 mixture-ratio sweep using one NASA9 chamber-equilibrium model and the explicitly selected frozen, equilibrium, or selected-area-freezing nozzle model. Chamber temperature is fixed or set by gas-phase energy closure.

MR = m_oxidizer/m_fuel; n_f = 1/M_f; n_ox = MR/M_ox; Each sample uses chamber equilibrium and the selected nozzle chemistry, with At, Ae/At, and pa held fixed

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - MR is oxidizer-to-fuel mass ratio. Every sample uses the same species data, chamber equilibrium, and Frozen from chamber exit nozzle, with fixed geometry and back pressure; it is not a lookup table interpolation across chemical assumptions. - Chamber temperature is fixed at the user-specified value for every sample; it is not a predicted adiabatic combustion temperature. - The optimal value is only the highest Isp among successfully sampled points, not a continuous or global optimum. Failure points remain and are not connected; no liquid vaporization, condensed phase, finite rate, or separation models.

History data: - mixture_ratio_sweep: 21 points; charts can select a quantity and JSON/CSV exports retain complete data.

Ramjet · Full flowpath · Subsonic combustion

Mode identifier: ramjet.full_engine.

Specified-recovery inlet, subsonic source-term flowpath, and nozzle at specified exit pressure; nozzle exit area is an output.

A specified-recovery inlet connects a subsonic source-term flow segment and a nozzle at specified exit static pressure. Exit area is recovered from mass continuity. This is not fixed-geometry back-pressure matching.

F = ṁe Ve − ṁair V₀ + (pe − p₀) Ae
S = ṁfuel/F

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Fixed specific heat, explicit heat flow, and mass sources; no reaction dynamics, shock trains, cooling coupling, or automatic mode switching. - Inlet total-pressure recovery is specified by the user. Capture area does not predict spillage or installation drag. - The subsonic combustor feeds a nozzle with specified exit static pressure. Exit area is calculated from mass continuity; this is not fixed-geometry back-pressure matching. - Only the nozzle inlet and exit states are reported. A line between them does not represent a resolved nozzle flowfield.

History data: - flow_profile: 201 points; charts can select a quantity and JSON/CSV exports retain complete data. - nozzle_stations: 2 points; charts can select a quantity and JSON/CSV exports retain complete data.

Ramjet · Full-engine design point

Mode identifier: ramjet.design_point.

Constant-cp relations. The quasi-one-dimensional flow model checks sonic events and integration residuals; it is not a complete inlet–combustor–nozzle chain.

Constant-cp relations. The quasi-one-dimensional flow model checks sonic events and integration residuals; it is not a complete inlet–combustor–nozzle chain. Subsonic-combustion ramjet design point, with no compressor or turbine shaft work. The combustor exit is closed at the specified total temperature; the ideal mode neglects fuel mass addition and requires complete expansion.

F = ṁe Ve − ṁair V₀ + (pe − p₀) Ae
S = ṁfuel/F

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The outlet is not fully expanded; the current kinetic energy decomposition does not include pressure thrust work, so no thermal or propulsive efficiency is reported. The total efficiency is still calculated based on thrust power/fuel chemical power.

Ramjet · Normal shock

Mode identifier: ramjet.normal_shock.

Constant-cp relations. The quasi-one-dimensional flow model checks sonic events and integration residuals; it is not a complete inlet–combustor–nozzle chain.

Calorically perfect-gas normal shock. Input is the upstream supersonic state; output includes downstream state ratios and total-pressure recovery.

M₂² = [1 + (γ − 1) M₁²/2]/[γ M₁² − (γ − 1)/2]

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Ramjet · Rayleigh flow · Constant-area heating

Mode identifier: ramjet.rayleigh.

Constant-cp relations. The quasi-one-dimensional flow model checks sonic events and integration residuals; it is not a complete inlet–combustor–nozzle chain.

Constant-area frictionless heat addition specified by total-temperature ratio. It retains the inlet Mach branch and checks the choking limit.

ρ u = constant
p + ρ u² = constant
Tt,2/Tt,1 = given value

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Ramjet · Oblique shock · Explicit weak/strong branch

Mode identifier: ramjet.oblique_shock.

Constant-cp relations. The quasi-one-dimensional flow model checks sonic events and integration residuals; it is not a complete inlet–combustor–nozzle chain.

Attached oblique shock with explicit weak or strong branch. The calculation stops beyond the attachment limit and does not substitute a normal shock.

tan θ = 2 cot β (M₁² sin²β − 1)/[M₁²(γ + cos 2β) + 2]

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Selected oblique shock branch: weak. The detachment shock wave is not calculated when the maximum attachment rotation angle is exceeded.

Ramjet · Fanno flow · Constant-area friction

Mode identifier: ramjet.fanno.

Constant-cp relations. The quasi-one-dimensional flow model checks sonic events and integration residuals; it is not a complete inlet–combustor–nozzle chain.

Constant-area adiabatic friction flow using the Darcy friction factor. It retains the inlet Mach branch and stops at the sonic limit.

Tt = constant
ṁ = ρ u A = constant
fD L/D = 4 fF L/D

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Adiabatic constant cross-section; using Darcy f_D·L/D (equal to 4 f_F·L/D), the total temperature remains constant.

Ramjet · Quasi-one-dimensional flow segment

Mode identifier: ramjet.one_d_flow.

Constant-cp relations. The quasi-one-dimensional flow model checks sonic events and integration residuals; it is not a complete inlet–combustor–nozzle chain.

Fixed-cp quasi-one-dimensional integration with area change, heat addition, mass addition, and friction. It checks sonic events and conservation residuals.

ṁ(x) = ρ(x) u(x) A(x)
M(x) = u(x)/√(γ R T(x))

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - With the specific heat fixed, the total enthalpy of injection uses the cpT zero point. - Uses a circular hydraulic diameter, Darcy friction, heat per unit length, and mass injection. This is not a complete reacting scramjet or isolator model.

History data: - flow_profile: 101 points; charts can select a quantity and JSON/CSV exports retain complete data.

Solid Rocket · Equilibrium Chamber Pressure

Mode identifier: solid_rocket.equilibrium.

Uses a supplied burn-rate law, burning area, and characteristic velocity. Demonstration parameters do not represent a specific propellant.

Uses a supplied burn-rate law, burning area, and characteristic velocity. Demonstration parameters do not represent a specific propellant. Zero-dimensional, constant-temperature combustion chamber with prescribed burn-rate law, characteristic velocity, and thrust coefficient. The transient calculation integrates grain regression and discharged mass, stopping at burnout or when it leaves the choked-flow domain. It does not model ignition, erosion, or tail-off blowdown.

r = rref (pc/pref)^n
ṁout = pc At/c*
mgas = mgas,0 + ρp Vburned − mout

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - burnArea is only used in equilibrium pressure mode; transients are calculated from the geometric burn surface.

History data: - history: 0 points; charts can select a quantity and JSON/CSV exports retain complete data.

Solid Rocket · Grain Regression Transient

Mode identifier: solid_rocket.transient.

Constant-temperature zero-dimensional chamber; end-burning or internal-port grain regression with conservative RK4 mass integration. No ignition, erosion, or tail-off.

Constant-temperature zero-dimensional chamber; end-burning or internal-port grain regression with conservative RK4 mass integration. No ignition, erosion, or tail-off. Zero-dimensional, constant-temperature combustion chamber with prescribed burn-rate law, characteristic velocity, and thrust coefficient. The transient calculation integrates grain regression and discharged mass, stopping at burnout or when it leaves the choked-flow domain. It does not model ignition, erosion, or tail-off blowdown.

r = rref (pc/pref)^n
ṁout = pc At/c*
mgas = mgas,0 + ρp Vburned − mout

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The specified time has been reached. - Fixed temperature lumped chamber pressure, given combustion rate law and thrust coefficient; excluding ignition, erosion combustion and nozzle geometry correction.

History data: - history: 2002 points; charts can select a quantity and JSON/CSV exports retain complete data.

TBCC · Explicit Air Split

Mode identifier: tbcc.explicit_split.

Combines independent turbojet and ramjet channels. Common inlet, back-pressure matching, and transition control are not implemented.

Combines independent turbojet and ramjet channels. Common inlet, back-pressure matching, and transition control are not implemented. Combines thrust and fuel flow from two independent cycles using a prescribed air split. It does not model a common inlet, back-pressure matching, or automatic mode switching.

F = F_turbojet + F_ramjet
ṁfuel = ṁfuel,turbojet + ṁfuel,ramjet

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - This result combines two independent cycles according to the air split ratio given by the user; the common inlet or back pressure matching is not solved. - This result does not include blending, variable geometry, channel gating, or automatic mode switching.

RBCC · Ideal Ejector-Ramjet Submodel

Mode identifier: rbcc.ideal_ejector.

Specified primary total state, supersonic primary flow, subsonic secondary flow, and choked ejector exit; not a complete RBCC.

Specified primary total state, supersonic primary flow, subsonic secondary flow, and choked ejector exit; not a complete RBCC. Ideal constant-area ejector-ramjet: a common gas constant, supersonic primary flow, subsonic secondary flow, choked ejector exit, and complete expansion. This is an RBCC submodel only.

α = ṁsecondary / ṁprimary
φ = (1+α) Vexit/Vprimary − α V₀/Vprimary

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The ideal model requires a supersonic primary-flow inlet, a subsonic secondary-flow inlet, and a choked ejector exit. - Does not include wall friction, shock waves, total pressure loss, fuel mass, secondary flow choke, or outlet unchoked branches. - The ideal ejection punch submodel only, cannot represent the complete RBCC or mode conversion.

ATR · Ideal Air-Turboramjet Submodel

Mode identifier: atr.ideal_air_turboramjet.

Specified primary flow, ideal work balance, and simplified heat addition; no gas generator, pump, or cooling.

Specified primary flow, ideal work balance, and simplified heat addition; no gas generator, pump, or cooling. Ideal air-turboramjet submodel with prescribed primary-stream total state: ideal fan/turbine work balance, low-speed mixing, simplified heat addition, and complete expansion. It excludes the gas generator, pumps, and cooling.

ṁprimary Δh_turbine = ṁair Δh_fan
F = ṁexit Vexit − ṁair V₀

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The prevailing total state is given by the input; the model does not solve for gas generators, pumps, cooling, or chemical equilibrium. - The ideal model assumes an isentropic fan and turbine, zero Mach mixing and combustion, no total pressure loss, and a fully expanding nozzle. - Exhaust branch: subsonic. The fuel ratio adopts a simplified heat addition formula, ignoring the correction of fuel consumption by the mainstream proportion and fuel outlet enthalpy; no accurate chemical heat balance prediction is made.

Scramjet · Calibrated-Inlet Full Engine · Nonreacting Mixed Flow

Mode identifier: scramjet.calibrated_mixing_engine.

Calibrated inlet–combustor–nozzle system. Given geometry and explicit injection define nonreacting mixing or finite-rate flow; captured flight momentum is subtracted. Ambient back-pressure matching is not automatic.

A calibrated dry-air inlet connects a given-geometry mixing or reacting-flow combustor and nozzle. Each segment may specify injection composition, total enthalpy, and velocity. Net thrust subtracts captured-air flight momentum. There is no fixed-geometry back-pressure matching or automatic transition; no full-engine thrust is reported when the solution is incomplete.

ṁair=ρ0V0Acapture；ṁout=ṁair+Σṁinj；Fnet=ṁoutVe+(pe−p0)Ae−ṁairV0

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The calibrated NASA9 inlet feeds nonreacting conservation-based mixed flow. Injection can change composition; reaction rates are zero and external heating is user-specified. - Net thrust subtracts captured-air momentum flux. specific_impulse_injected uses total injected mass; injected oxidizer or air is not relabeled as fuel. - User-specified reaction coefficients and preset dissociation demonstrations are not a substitute for confirmed combustion mechanisms. - Downstream geometry is fixed by the input. Shocks, isolator states, and area are not adjusted to match ambient back pressure. Spillage drag and automatic mode transition are excluded. - If any segment is incomplete, only its valid prefix is exported and full-engine thrust is not reported.

History data: - full_flow_profile: 140 points; charts can select a quantity and JSON/CSV exports retain complete data.

Scramjet · Calibrated-Inlet Full Engine · Finite-Rate Reacting Flow

Mode identifier: scramjet.calibrated_reacting_engine.

Calibrated inlet–combustor–nozzle system. Given geometry and explicit injection define nonreacting mixing or finite-rate flow; captured flight momentum is subtracted. Ambient back-pressure matching is not automatic.

A calibrated dry-air inlet connects a given-geometry mixing or reacting-flow combustor and nozzle. Each segment may specify injection composition, total enthalpy, and velocity. Net thrust subtracts captured-air flight momentum. There is no fixed-geometry back-pressure matching or automatic transition; no full-engine thrust is reported when the solution is incomplete.

ṁair=ρ0V0Acapture；ṁout=ṁair+Σṁinj；Fnet=ṁoutVe+(pe−p0)Ae−ṁairV0

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - The calibrated NASA9 dry-air inlet feeds a combustor and nozzle using user-specified finite-rate reactions. Injection composition and formation-inclusive enthalpy are specified per segment. - Net thrust subtracts captured-air momentum flux. specific_impulse_injected uses total injected mass; injected oxidizer or air is not relabeled as fuel. - User-specified reaction coefficients and preset dissociation demonstrations are not a substitute for confirmed combustion mechanisms. - Downstream geometry is fixed by the input. Shocks, isolator states, and area are not adjusted to match ambient back pressure. Spillage drag and automatic mode transition are excluded. - If any segment is incomplete, only its valid prefix is exported and full-engine thrust is not reported.

History data: - full_flow_profile: 437 points; charts can select a quantity and JSON/CSV exports retain complete data.

Scramjet · Finite-Rate · Quasi-One-Dimensional Reacting Flow

Mode identifier: scramjet.finite_rate_flow.

User-supplied finite-rate reactions coupled to quasi-one-dimensional mass, momentum, and formation-enthalpy energy equations. Explicit Mach branch; third-body and Troe support; no automatic transonic transition.

Mass-action rates close with species mass flow, F/G, and NASA9 formation enthalpy, so temperature is not fixed. The inlet Mach branch is explicit; geometry, friction, and heat source may be specified; injection composition is set per segment or inherited from the inlet. Supports third-body and Lindemann/Troe chemistry, but not transport or transonic matching.

dNᵢ/dx = Aωᵢ + ṁ′Yᵢ,inj/Wᵢ；dF/dx = pA′ − τP + ṁ′Vinj；d(ṁG)/dx = q′ + ṁ′Ginj

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Quasi-one-dimensional finite rate reaction flow: temperature is closed with the enthalpy of formation and total energy; no additional LHV is added to the reaction heat release. F/G reconstruction retains the selected inlet Mach branch. - Reaction coefficients are user inputs; the default dissociation reaction demonstrates the algorithm only. Supports third-body effects and Lindemann/Troe falloff. Excludes diffusion, shock trains, and automatic transonic matching. - Injection composition can be specified per segment; otherwise inlet composition is used. Injection total enthalpy includes formation enthalpy; explicit heat input represents additional external heating. A valid prefix from an incomplete integration is not a complete nozzle or full-engine result. - Thermal data: NASA9 · 13 species, version: NASA9-c915968268c707d69776909cadd834f1bcb6e44e4aaf1adfc6d5f42f86f6ace4.

History data: - reacting_flow_profile: 21 points; charts can select a quantity and JSON/CSV exports retain complete data.

Scramjet · Seven-State Variable-Specific-Heat · Frozen Flow Segment

Mode identifier: scramjet.variable_cp_flow.

NASA9 seven-state/F–G branch reconstruction and same-composition frozen flow with formation enthalpy in heat and mass injection; no chemical reaction.

NASA9 seven-state inlet reconstruction with explicit subsonic or supersonic F/G roots; G includes formation enthalpy. Flow segments allow area change, heat input, same-composition injection, and friction. Mass and energy residuals are checked; sound-speed or property-range events stop the solution. No along-path chemical reaction.

F = pA + ṁV；G = h(T,x) + V²/2；p = ρRmixT；d(ṁG)/dx = q′ + ṁ′Ginj

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - NASA9 seven-state reconstruction and frozen composition flow segment; G=h+V²/2 includes the enthalpy of formation, allowing physically valid negative values. - The injection stream must have the same composition as the main stream; the explicit heat source does not equal the reaction mechanism, and the mixing or chemical rates of different components are not calculated.

History data: - frozen_flow_profile: 201 points; charts can select a quantity and JSON/CSV exports retain complete data.

Scramjet · Reference Calibration · Variable-Specific-Heat Inlet

Mode identifier: scramjet.calibrated_variable_inlet.

NASA9 dry air with explicit reference entropy-rise calibration, three freestream definitions, five inlet closures, and multiple-root selection; an independent inlet model.

A reference point explicitly calibrates Δs/ln(pe/p0). Three freestream definitions and five inlet closures share total-enthalpy, mass, and entropy relations. Uses NASA9 dry air from 0–70 km. Contraction inversion explicitly selects the higher- or lower-Mach root; it is not matched to a complete reacting flowpath.

h + V²/2 = h0；CR = ρeVe/(ρ0V0)；Δs = K ln(pe/p0)；πt = exp(−Δs/R)

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Reference calibration uses Δs = K ln(pe/p0), NASA9 dry-air properties with frozen composition, and total-enthalpy conservation. It does not substitute local γ into a constant-γ total-pressure power law. - Reference altitude, flight Mach, exit Mach, and contraction ratio are saved explicitly. High/low Mach branches apply to contraction-ratio inversion only; roots near extrema are rejected when they cannot be bracketed. - Three freestream specifications and five exit constraints are supported over 0–70 km geometric altitude. Calculations use the bundled NASA9 data; numerical equivalence to legacy property implementations is not claimed. - Independent inlet model; it does not by itself match a variable-specific-heat reacting combustor or fixed-geometry ambient back pressure.

Scramjet · Finite-Rate · Constant-Temperature, Constant-Volume Reactor

Mode identifier: scramjet.finite_rate_reactor.

User-supplied mass-action reactions in a constant-volume, externally isothermal reactor with adaptive explicit integration. Supports third-body/Troe and an optional NASA7 hydrogen–oxygen submechanism; no transport.

Constant-volume, externally isothermal mass-action reactor. It can load the 29-reaction hydrogen–oxygen submechanism or custom coefficients, supports third-body/Lindemann/Troe effects, and derives reverse rates from selected thermodynamic data. The initial single dissociation is numerical demonstration only. Incomplete integration cannot be used for extrema analysis.

dcᵢ/dt = Σ νᵢⱼ (kf ∏cᵣ^νᵣ − kr ∏cₚ^νₚ)；kf = A Tⁿ exp(−Ea/RuT)；kr = kf/Kc

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Constant volume, external constant temperature reactor; temperature is not updated by heat release, thermostat_heat_rate is the heat flow per unit volume required to maintain constant temperature. - Reaction-rate coefficients are supplied by the user. The default H2→2H reaction with A=1 s⁻¹ demonstrates the algorithm; it is not a calibrated hydrogen-combustion mechanism. - Arrhenius A uses concentrations in kmol/m³ and temperature in K; Ea is in J/kmol. Reverse rates use equilibrium constants derived from the selected Gibbs-energy data. Third-body effects and Lindemann/Troe falloff are supported; transport is excluded. - Thermal data: NASA9 · 13 species, version: NASA9-c915968268c707d69776909cadd834f1bcb6e44e4aaf1adfc6d5f42f86f6ace4.

History data: - reaction_history: 178 points; charts can select a quantity and JSON/CSV exports retain complete data.

Scramjet · Specified Temperature and Pressure · Gas-Phase Equilibrium

Mode identifier: scramjet.gas_equilibrium_tp.

Fixed T/p, 13-species ideal-gas equilibrium with explicit elemental inventory and chemical-affinity residuals; no ions or condensed phase.

At fixed temperature and pressure, Gibbs energy is minimized for the listed 13 ideal-gas species while conserving elements. Unconverged results are not reported. This mode does not solve adiabatic flame temperature.

Σ νᵢ [gᵢ°/(RuT) + ln(xᵢ p/p°)] = 0；Σ aₖᵢ nᵢ = bₖ

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - Ideal gas-phase reaction equilibrium for a given T, p; uses a 13-species ensemble and does not contain ionic, condensate, or complete hydrocarbon mechanisms. - Reported cp and γ are frozen-composition properties evaluated at the equilibrium composition, not derivatives along a changing chemical-equilibrium path. - Initial species amounts define the elemental inventory only; this mode does not solve adiabatic combustion temperature. Data version: NASA9-c915968268c707d69776909cadd834f1bcb6e44e4aaf1adfc6d5f42f86f6ace4.

Scramjet · Constant-Pressure Energy Closure · Gas-Phase Equilibrium

Mode identifier: scramjet.gas_equilibrium_hp.

Inlet gaseous composition, temperature, pressure, and heat input close equilibrium enthalpy. Zero heat is adiabatic. No liquid or condensed phase.

At fixed pressure and elemental inventory, inlet gas-phase enthalpy and specified heat determine equilibrium temperature and composition. Every temperature trial must converge and the root interval is explicit. No liquid phase, condensed phase, or ions.

h_eq(T,p,b) = h_in(T_in,x_in) + q; Σ aₖᵢ nᵢ = bₖ; q = 0 is adiabatic

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - At specified pressure, the inlet gaseous composition and common inlet temperature define the formation-inclusive enthalpy used for energy closure. Zero heat input gives adiabatic gas-phase equilibrium. - Only the 13 species declared in the gas phase; no liquid vaporization, condensed phase, ions, reaction times or combustion efficiencies. Output is rejected for solutions outside the interval and that have not converged. - Reported cp and γ are frozen-composition properties of the resulting equilibrium mixture. NASA9 data version: NASA9-c915968268c707d69776909cadd834f1bcb6e44e4aaf1adfc6d5f42f86f6ace4.

Scramjet · Variable-Specific-Heat · Frozen Nozzle

Mode identifier: scramjet.variable_cp_frozen_nozzle.

NASA9 variable-property frozen composition, isentropic enthalpy drop, and supersonic area matching; not replaced by constant-γ power laws.

NASA9 variable-property nozzle with frozen composition. Mass, total enthalpy, and entropy determine area; the supersonic branch does not cross sonic conditions. Extrapolation outside the property range is rejected.

h(T) + V²/2 = h₀；s(T,p) = s₀；ρVA = ṁ；Fgross = ṁVe + (pe − p₀) Ae

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - NASA9 variable specific heat with frozen composition and adiabatic, isentropic supersonic expansion. Excludes reactions, friction, separation, and side loads. - gross_thrust is nozzle gross thrust. Full-engine net thrust also subtracts captured-air momentum flux. - Physical property data version: NASA9-c915968268c707d69776909cadd834f1bcb6e44e4aaf1adfc6d5f42f86f6ace4.

History data: - nozzle_profile: 101 points; charts can select a quantity and JSON/CSV exports retain complete data.

Scramjet · Frozen Composition · Variable-Specific-Heat Properties

Mode identifier: scramjet.frozen_gas_properties.

NASA9 frozen ideal mixture with variable heat capacity, formation enthalpy, mixing entropy, and enthalpy inversion; no chemical equilibrium.

NASA9 frozen ideal mixture with explicit formation-enthalpy reference and mixing entropy. It is calculated only within the common temperature range of all selected species and is not chemical equilibrium.

h = Σ xᵢ hᵢ / Mmix；cp = Σ xᵢ cp,ᵢ / Mmix；cv = cp − Rmix；γ = cp/cv

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.

Model notes for this demonstration: - NASA9 ideal-gas properties with frozen composition, including formation enthalpy and ideal mixing entropy; chemical equilibrium is not solved. - Physical property data version: NASA9-c915968268c707d69776909cadd834f1bcb6e44e4aaf1adfc6d5f42f86f6ace4. - Enthalpy includes formation enthalpy and must not be mixed with a cp·T reference convention. Temperature extrapolation is not supported.

Scramjet · Full Engine · Constant-Specific-Heat Flowpath

Mode identifier: scramjet.full_engine.

Seven-state input, fixed-cp explicit heat and mass sources, and a frozen nozzle; no reaction kinetics.

Seven-state input, fixed-cp explicit heat and mass sources, and a frozen nozzle; no reaction kinetics. Constant-specific-heat inlet, prescribed source-term flow segments, and a frozen nozzle. F is the momentum function pA + ṁV and G is total enthalpy per unit mass. Integration stops at sonic events and does not cross singularities.

F = p A + ṁ V
G = cp T + V²/2
Fnet = ṁe Ve − ṁair V₀ + (pe − p₀) Ae

Workflow: select this engine and mode in the app, choose Load Current Mode Demonstration, adjust fields, then calculate. The table lists numeric fields visible for the demonstration configuration; the app presents the current full field set after optional-field, freestream, or component changes. All table values are SI.