CAD Models & Simulations Gallery

Discover a curated collection of high-fidelity CAD models, showcasing rocket engines, turbopumps, aerospace flow instrumentation, and re-entry vehicles. Each model is supported by detailed engineering analysis, including CFD simulations of shockwaves, hypersonic flow around re-entry bodies, and thermal and structural evaluations under extreme operating conditions. From detached shock wave visualization in high-Mach atmospheric re-entry to turbopump flow dynamics and thrust chamber cooling, every project features 3D visualizations, performance plots, and technical documentation—highlighting real-world design intent and functional behavior.

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Rocket Engine Turbopump

A detailed model of a rocket engine turbopump assembly. This critical component is responsible for delivering propellants at high pressure and flow rates to the combustion chamber. The design features both turbine and pump sections with precision-engineered impellers and housing.

Technical Analysis

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CFD
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CFD

Design Features

    The turbopump assembly is designed with a strategic choice of materials to balance structural performance, chemical compatibility, and mass efficiency. The volute casing is 6061-T6 aluminum alloy fabricated from , selected for its lightweight properties, good corrosion resistance, and ease of machining—ideal for non-rotating, non-direct-contact regions where reducing system mass is critical. In contrast, the inducer, impeller, and stator vanes are made from Ti-6Al-4V titanium alloy, owing to its high specific strength, excellent fatigue resistance, and compatibility with cryogenic oxidizers such as liquid oxygen. Titanium also offers superior erosion resistance in high-speed rotating components and maintains mechanical integrity under extreme pressure and thermal gradients. This material configuration ensures a robust and efficient turbopump capable of operating reliably under demanding rocket propulsion conditions.
  • Staging: Multi-stage centrifugal pump, with first stage inducer and second stage impeller
  • Fluid: RP-1
  • Mass flow rate: 100kg/s
  • Speed: 35,000 RPM
  • Pressure difference: 175 bar
  • Specific speed: 38.3
  • Specific work: 21,472 m^2/s^2
  • Power output: 2147.2 kW
  • Head: 2189.6 m

Compressor for Turboshaft engine

A detailed model of a compressor for turboshaft engine, used in helicopters.

Design Features

  • Staging: Single Stage
  • Fluid: Air
  • Mass flow rate: 4kg/s
  • Speed: 26,000 RPM
  • Pressure difference: 1.5 bar
  • Specific speed: 100.8
  • Specific work: 36,181 m^2/s^2
  • Power input: 144.72 kW

Showerhead Injector Design

A Showerhead Injector is a type of propellant injector commonly used in liquid rocket engines. It consists of multiple small orifices arranged in a showerhead-like pattern, typically on a flat faceplate. Each orifice directs fuel and oxidizer into the combustion chamber, promoting atomization and mixing. This injector is known for its simplicity and ease of manufacturing. It is especially suitable for small-scale or experimental engines and is often used in educational or prototype rocket systems.

Merits :
• Simple Design - Easy to fabricate with standard machining tools.
• Cost-Effective - Lower manufacturing and development costs.
• Uniform Distribution - Provides consistent propellant distribution across the chamber face.
• Reliable - Fewer components reduce the risk of mechanical failure.

Demerits :
• Poor Mixing Efficiency - Compared to swirl or impinging injectors, the mixing quality is lower.
• Limited Atomization - Droplet breakup and vaporization are less efficient.
• Risk of Combustion Instabilities - May lead to uneven combustion and performance fluctuations.
• Lower Performance - Not ideal for high-performance or high-thrust applications.

Technical Specifications

  • Total mass flow rate: 513.16231 kg/s
  • Oxidizer mass flow rate: 440.16625 kg/s
  • Fuel mass flow rate: 72.99606 kg/s
  • Number of orifice for fuel flow: 500
  • Diameter of each fuel orifice: 1.944 mm
  • Number of orifice for oxidizer flow: 1000
  • Diameter of each oxidizer orifice: 3.12 mm

Coaxial Swirl Injector Design

A coaxial swirl injector is a type of liquid rocket engine injector designed to achieve efficient atomization and mixing of propellants. In this configuration, one propellant (usually oxidizer) flows through a central orifice while the other (usually fuel) is introduced tangentially through swirl passages, creating a hollow conical spray sheet. The high shear between the coaxial streams breaks the liquid film into fine droplets, promoting rapid evaporation and combustion stability.

Merits :
• Produces very fine droplets, enhancing mixing and combustion efficiency.
• Capable of stable operation across a wide range of flow rates.
• Helps reduce combustion instabilities due to uniform mixing.
• Compact design, suitable for high-performance engines.

Demerits :
• More complex to manufacture compared to simple orifice injectors.
• Swirl passages are prone to clogging or erosion under long-term operation.
• Higher pressure drop may be required to maintain effective atomization.

Technical Analysis

CFD
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CFD

Ionic Thruster Mark 1

This model represents a compact ion propulsion system, designed for long-duration space missions where high efficiency and precision thrust control are essential. The thruster operates by ionizing xenon gas, accelerating the resulting ions through an electrostatic field, and expelling them to generate thrust with extremely high specific impulse (Isp).

Technical Specifications

  • Electron Inlet (Discharge Cathode): The large-diameter pipe feeds electrons directly into the chamber to ionize the xenon atoms. These electrons are emitted from a cathode (typically a hollow cathode or thermionic emitter), and collide with neutral xenon atoms to generate positively charged ions (Xe⁺) through impact ionization.
  • Xenon Inlet System: A series of small-diameter inlet tubes, arranged in a ring around the central chamber, introduce xenon gas uniformly into the ionization region. This ensures a stable and symmetric plasma discharge, vital for consistent thrust.
  • Grid System: At the thruster's exit, a finely spaced grid structure is visible. This forms the acceleration stage, where high-voltage electrodes create an electrostatic field to accelerate positively charged xenon ions. The grid ensures beam collimation and minimizes ion divergence, improving propulsion efficiency.
  • Electron Source (Neutralizer, at the end): The pipe at the end is used to inject electrons into the exhaust plume. These electrons neutralize the positive ion beam, preventing spacecraft charge buildup and maintaining electric balance in space.

Ionic Thruster Mark 2

This is a CAD model of a compact ion thruster, designed for deep space missions and low-Earth orbit satellite station-keeping. The design features a cylindrical discharge chamber with a central ionization zone, enclosed between structural support plates and integrated with electrical feedthroughs for grid or cathode connections.
Key design aspects:

Central ion emitter or cathode, potentially coupled with a gas feed (e.g., Xenon or Argon) to generate and accelerate ions.

Mounting rods and brackets to ensure rigidity and easy CubeSat integration.

Highly reflective metallic finish, optimized for thermal resistance and space-grade durability.

This thruster aims to deliver low thrust with high specific impulse, making it ideal for applications such as deep space navigation, constellation orbit maintenance, and attitude control. Future work includes plasma plume simulation, ion beam divergence analysis, and integration with satellite power systems.

Technical Specifications

    The thruster is constructed using material-specific design zones to ensure durability and performance in the harsh environment of space. Components that come into direct contact with the ionized plasma, such as the central discharge nozzle and inner acceleration surfaces, are fabricated from titanium alloy (shown in yellow-bronze color). Titanium offers excellent corrosion resistance, low sputter yield, and high thermal tolerance, making it ideal for sustained ion exposure. In contrast, the structural frame, support rods, and mounting plates are made from 6061-T6 aluminum alloy, chosen for its lightweight properties, high strength-to-weight ratio, and ease of machining—suitable for areas without direct ion impact. This strategic material segregation ensures both mission durability and mass efficiency, critical for satellite propulsion systems.

Supersonic flow over a double wedge

This project analyzes supersonic flow at Mach 3 over a double wedge geometry to study shock wave interactions and high-speed aerodynamic behavior. The sharp angles generate oblique shocks and complex shock-shock interactions, which are critical in the design of high-speed aerospace vehicles. The simulation captures pressure distribution, Mach contours, and flow separation zones, offering insights into wave dynamics relevant to supersonic and hypersonic applications.

Technical Analysis

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2D Engineering Sketch with Dimensions
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Velocity Magnitude
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Absolute Pressure
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Density
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Enthalpy
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Temperature

Technical Specifications

  • Boundary Conditions: Walls and pressure farfield
  • Velocity: Mach 3

Space Vehicle Re-entry CFD Simulation

This project involves the Computational Fluid Dynamics (CFD) simulation of a space vehicle during atmospheric re-entry, conducted using ANSYS Fluent. The simulation captures critical hypersonic flow phenomena such as aerodynamic heating. The analysis focused on temperature and pressure variation around the vehicle surface. High-temperature gradients at the stagnation point and along the heat shield, essential for Thermal Protection System (TPS) design.

Technical Analysis

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Model
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2D Temperature Contour
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Static Pressure
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Temperature
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Temperature

Technical Specifications

  • Boundary Conditions: Walls and pressure farfield
  • Velocity: 3D simulation run at Mach 5 and 2D at Mach 24

Detached Shock Wave CFD Simulation over a Sharp-Nosed Body

This simulation shows a detached shock wave forming ahead of a sharp-nosed re-entry body in hypersonic flow, modeled using ANSYS Fluent. Key observations: A detached bow shock, typically unexpected for sharp geometries, occurs here due to specific flow conditions such as high Mach number and low Reynolds number or thermal effects. Intense aerodynamic heating at the nose region, visible through the high-temperature contour (red). Gradual pressure and temperature gradients along the surface, relevant for evaluating thermal protection system (TPS) design. This case demonstrates that even sharp geometries can produce detached shocks in certain high-enthalpy flows, making it a valuable study in advanced hypersonic vehicle design and aerothermal analysis.

Technical Analysis

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Model
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Static Pressure
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Temperature
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Velocity

Technical Specifications

  • Boundary Conditions: Walls and pressure farfield
  • Velocity: Mach 1.5

RS-25 Rocket Engine

A detailed 3D model of the RS-25 (Space Shuttle Main Engine), one of the most advanced rocket engines ever built. This model showcases the complex internal structure and engineering excellence of this liquid-fueled rocket engine.

Technical Analysis

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2D Engineering Sketch with Dimensions
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Cross-sectional View
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CFD

Technical Specifications

  • Thrust: 2,279 kN at vacuum
  • Specific Impulse: 452.3 seconds (vacuum)
  • Propellants: Liquid Hydrogen / Liquid Oxygen
  • Chamber Pressure: 20.64 MPa
  • Mixture Ratio: 6.03:1 (O2:H2)
  • Total mass flow rate: 513.16231 kg/s
  • Oxidizer mass flow rate: 440.16625 kg/s
  • Fuel mass flow rate: 72.99606 kg/s

Chamber Performance Data

Parameter Injector Nozzle Inlet Nozzle Throat Nozzle Exit Unit
Pressure 20.6400 20.1170 11.7132 0.0181 MPa
Temperature 3603.8962 3597.4647 3386.6714 1202.3447 K
Enthalpy -983.8302 -1012.2874 -2156.0939 -10653.1087 kJ/kg
Entropy 17.1304 17.1382 17.1382 17.1382 kJ/(kg·K)
Internal energy -3177.0249 -3201.1430 -4198.2609 -11358.5321 kJ/kg
Specific heat (p=const) 7.3583 7.3598 6.7559 2.8775 kJ/(kg·K)
Specific heat (V=const) 6.2932 6.2952 5.8003 2.2908 kJ/(kg·K)
Gamma 1.1693 1.1691 1.1648 1.2561 -
Isentropic exponent 1.1470 1.1470 1.1481 1.2561 -
Gas constant 0.6086 0.6085 0.6030 0.5867 kJ/(kg·K)
Molecular weight (M) 13.6625 13.6639 13.7885 14.1716 -
Molecular weight (MW) 0.01366 0.01366 0.01379 0.01417 -
Density 9.4109 9.1898 5.7357 0.0256 kg/m³
Sonic velocity 1586.0840 1584.5308 1531.1847 941.3175 m/s
Velocity 0.0000 238.5673 1531.1847 4397.5626 m/s
Mach number 0.0000 0.1506 1.0000 4.6717 -
Area ratio 4.0000 4.0000 1.0000 78.0000 -
Mass flux 2192.3834 2192.3834 8782.3673 112.6356 kg/(m²·s)
Mass flux (relative) 1.0626 1.0904 - - kg/(N·s)
Viscosity 0.0001085 0.0001084 0.0001037 4.493e-05 kg/(m·s)
Conductivity, frozen 0.5765 0.5756 0.5425 0.184 W/(m·K)
Specific heat (p=const), frozen 3.786 3.785 3.75 2.877 kJ/(kg·K)
Prandtl number, frozen 0.7712 0.7124 0.7167 0.7027 -
Conductivity, effective 1.429 1.428 1.247 0.184 W/(m·K)
Specific heat (p=const), effective 7.358 7.36 6.756 2.877 kJ/(kg·K)

Species Composition Data

Species Injector
mass fractions		 Injector
mole fractions		 Nozzle inlet
mass fractions		 Nozzle inlet
mole fractions		 Nozzle throat
mass fractions		 Nozzle throat
mole fractions		 Nozzle exit
mass fractions		 Nozzle exit
mole fractions
H 0.0018867 0.0255735 0.0018845 0.0255464 0.0015061 0.0206033 0.0000002
H2 0.0360445 0.2442888 0.0360324 0.2442322 0.0353322 0.2416701 0.0341731 0.2402366
H2O 0.9074897 0.6882250 0.9077116 0.6884641 0.9237487 0.7070161 0.9658269 0.7597632
H2O2 0.0000430 0.0000176 0.0000430 0.0000173 0.0000212 0.0000086 - -
HO2 0.0000873 0.0000361 0.0000859 0.0000355 0.0000410 0.0000171 - -
O 0.0024718 0.0021107 0.0024613 0.0021020 0.0015069 0.0012987 - -
O2 0.0053295 0.0022755 0.0053116 0.0022681 0.0034103 0.0014695 - -
OH 0.0466465 0.0374725 0.0464696 0.0373342 0.0344334 0.0279165 - -

Throttled Chamber Performance

Throttle value Pressure, MPa c ef (SL), m/s Is (SL), s Thrust (SL), kN c ef (opt), m/s Is (opt), s Thrust (opt), kN c ef (vac), m/s Is (vac), s Thrust (vac), kN
0.6700 13.79 3290.838 335.572 1131.453 4269.282 435.346 1467.860 4430.645 451.800 1523.340
0.7120 14.66 3329.880 339.553 1216.644 4272.618 435.686 1561.094 4433.829 452.125 1619.996
0.7540 15.53 3365.468 343.182 1302.182 4275.754 436.056 1654.394 4436.325 452.330 1716.717
0.7960 16.40 3397.900 346.489 1387.965 4278.518 436.287 1747.678 4439.458 452.699 1813.418
0.8380 17.28 3425.669 349.321 1473.141 4279.017 436.338 1840.107 4439.835 452.737 1909.263
0.8800 18.15 3451.044 351.990 1558.432 4279.366 436.386 1932.543 4440.189 452.773 2005.113
0.9220 19.02 3474.989 354.350 1644.141 4279.926 436.431 2024.986 4440.521 452.807 2100.969
0.9409 19.41 3485.175 355.389 1682.825 4280.117 436.450 2066.664 4440.665 452.822 2144.185
0.9640 19.89 3507.654 357.681 1735.196 4280.342 436.473 2117.436 4440.835 452.839 2196.830
1.0000 20.64 3541.503 361.133 1819.638 4280.496 436.508 2196.684 4441.032 452.865 2279.000
1.0260 20.76 3546.909 361.684 1831.061 4280.735 436.513 2209.892 4441.131 452.869 2292.695
1.0480 21.64 3583.027 365.367 1926.931 4281.107 436.551 2302.354 4441.412 452.898 2388.565
1.0900 22.51 3616.368 368.767 2022.804 4281.460 436.587 2394.821 4441.679 452.925 2484.439

CE-20 Rocket Engine

The CE-20 is an Indian cryogenic rocket engine developed by ISRO. This upper stage engine uses liquid hydrogen and liquid oxygen propellants and represents a significant achievement in Indian space technology.

Technical Analysis

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CFD

Technical Specifications

  • Thrust: 186.36 kN (41,900 lbf) in vacuum
  • Specific Impulse: 443 seconds (vacuum)
  • Propellants: Liquid Hydrogen / Liquid Oxygen
  • Chamber Pressure: 5.88 MPa (853 psi)
  • Burn Time: 640 seconds

Chamber Performance Data

Parameter Injector Nozzle inlet Nozzle throat Nozzle exit Unit
Pressure 6.0000 5.8470 3.3895 0.0033 MPa
Temperature 3285.1914 3279.1233 3068.4545 858.0217 K
Enthalpy -1077.4925 -1107.4919 -2312.2874 -11068.9544 kJ/kg
Entropy 19.4088 19.4178 19.4178 19.4178 kJ/(kg·K)
Internal energy -3374.3023 -3399.8056 -4440.5398 -11653.8965 kJ/kg
Specific heat (p=const) 7.3136 7.3086 6.5122 2.9463 kJ/(kg·K)
Specific heat (V=const) 6.2479 6.2442 5.5627 2.2645 kJ/(kg·K)
Gamma 1.1706 1.1705 1.1707 1.3010
Isentropic exponent 1.1554 1.1554 1.1556 1.1604
Gas constant 0.6991 0.6991 0.6996 0.6817 kJ/(kg·K)
Molecular weight (M) 11.8924 11.8937 11.9876 12.1961
Molecular weight (MW) 0.01189 0.01189 0.01199 0.0122
Density 2.6123 2.5507 1.5926 0.0056 kg/m³
Sonic velocity 1629.0565 1627.4443 1571.4917 872.3741 m/s
Velocity 0.0000 244.4967 1571.4917 4470.2264 m/s
Mach number 0.0000 0.1505 1.0000 5.1242
Area ratio 4.0000 4.0000 1.0000 100.0000
Mass flux 624.7807 624.7807 2502.7763 25.0198 kg/(m²·s)
Mass flux (relative) 1.041e-04 1.069e-04 kg/(N·s)
Viscosity 9.845e-05 9.831e-05 9.361e-05 3.254e-05 kg/(m·s)
Conductivity, frozen 0.5968 0.5958 0.5528 W/(m·K)
Specific heat (p=const), frozen 4.136 4.135 4.028 2.946 kJ/(kg·K)
Prandtl number, frozen 0.6824 0.6824 0.6887 0.6055
Conductivity, effective 0.5968 0.5958 0.5528 0.1583 W/(m·K)
Specific heat (p=const), effective 7.314 7.309 6.512 2.946 kJ/(kg·K)
Prandtl number, effective 0.5057 0.5055 0.5134 0.6055

Species Composition Data

Species Injector mass fractions Injector mole fractions Nozzle inlet mass fractions Nozzle inlet mole fractions Nozzle throat mass fractions Nozzle throat mole fractions Nozzle exit mass fractions Nozzle exit mole fractions
H 0.0022903 0.0270224 0.0022839 0.0269498 0.0016715 0.0198794 0.0016974 0.0198794
H2 0.0598531 0.3530959 0.0598464 0.3530954 0.0596967 0.3549907 0.0601180 0.3637140
H2O 0.9066409 0.5985006 0.9068513 0.5987058 0.9189761 0.6114971 0.9398820 0.6362860
H2O2 0.0000081 0.0000028 0.0000079 0.0000028 0.0000032 0.0000011
HO2 0.0000153 0.0000055 0.0000150 0.0000054 0.0000055 0.0000020
O 0.0011307 0.0008404 0.0011212 0.0008335 0.0005413 0.0004056
O2 0.0014905 0.0005540 0.0014790 0.0005497 0.0007343 0.0002751
OH 0.0285710 0.0199784 0.0283952 0.0198576 0.0183713 0.0129489

Throttled Chamber Performance

Throttle value Pressure, MPa c ef (SL), m/s Isp (SL), s Thrust (SL), kN c ef (opt), m/s Isp (opt), s Thrust (opt), kN c ef (vac), m/s Isp (vac), s Thrust (vac), kN
0.9658 5.79 2374.690 242.151 96.114 4314.308 439.937 174.619 4445.210 453.285 179.917
0.9914 5.95 2409.963 245.748 100.131 4315.631 440.072 179.310 4446.496 453.416 184.747
1.0000 6.00 2421.385 246.913 101.475 4316.065 440.116 180.876 4446.918 453.459 186.360
1.0171 6.10 2443.670 249.185 104.158 4316.919 440.203 184.002 4447.748 453.544 189.578
1.0427 6.26 2475.921 252.474 108.193 4318.175 440.331 188.695 4448.969 453.669 194.411
1.0684 6.41 2506.814 255.624 112.368 4319.399 440.456 193.390 4450.159 453.790 199.245
1.0940 6.57 2536.439 258.645 116.288 4320.593 440.578 198.086 4451.320 453.908 204.080
1.1196 6.72 2564.877 261.545 120.348 4321.758 440.697 202.783 4452.453 454.024 208.916
1.1453 6.88 2592.044 264.315 124.408 4322.928 440.816 207.483 4453.584 454.139 213.754
1.1709 7.03 2618.331 266.995 128.283 4324.040 440.929 212.183 4454.666 454.249 218.593
1.1966 7.19 2643.637 269.576 132.565 4325.127 441.040 216.884 4455.724 454.357 223.432
1.2222 7.34 2668.021 272.062 136.655 4326.191 441.149 221.585 4456.759 454.463 228.273