Neo Materials / Services / Aerospace

Aerospace & Spacecraft

Structural FEA · Orbital Mechanics · Propulsion · Composite Structures · MIL-SPEC Design

From hypersonic re-entry vehicles to cubesat structures, Neo Materials delivers aerospace engineering grounded in aeroelasticity, orbital mechanics, thermal protection systems, and composite structural design. Every design is FEA-verified, weight-optimised, and mission-specific — from 1U cubesats to 200T launch vehicle elements.

±0.01 mm
Structure Tol.
FEA Verified
ANSYS/Nastran
ECSS/NASA
Standards
MIL-SPEC
Compliant
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AEROSPACE ENGINEERING DIVISION

Aerospace & Spacecraft: Full Capability Tour

10 discipline modules covering every aspect of aerospace structural and systems engineering.

Aerospace Structural Engineering
Primary structures designed to survive combined axial, bending, torsion, and vibration loads across mission lifetime. Structural design follows ECSS-E-ST-32 (ESA) and NASA-STD-5001B allowables with minimum safety factor 1.25 on yield and 1.4 on ultimate.
Safety factor (yield)1.25 minimum
Safety factor (ultimate)1.40 minimum
Primary standardECSS-E-ST-32 / NASA-STD-5001
Load casesLaunch + Maneuver + Landing
Structural analysisANSYS Mechanical / MSC Nastran
Weight efficiencyTopology optimised (SIMP)

Topology optimisation using SIMP (Solid Isotropic Material with Penalisation) reduces structural mass by 30–50% while maintaining stiffness. Applied to launch vehicle brackets, satellite primary structure, and fairings.

LAUNCH VEHICLE PRIMARY STRUCTURE Ring frame Stringers σ_θ (hoop) F_axial P_int Material: Al-2024-T3 monocoque Wall thickness: t = 2.5 mm optimised Critical load: N_cr = π²EI/L² = 850 kN Buckling SF = 1.8 (knockdown 0.65 applied)
Aerodynamic Analysis & CFD
Compressible flow analysis from subsonic to hypersonic regimes (M 0.1–25). CFD using RANS (k-ω SST turbulence model) for transonic drag, panel methods for subsonic lift, and Newtonian flow for hypersonic pressure coefficients. Aeroelastic coupling with structural modes.
CFD solverOpenFOAM / ANSYS Fluent
Turbulence modelk-ω SST (transitional)
Mach range0.1 – 25 (hyp.)
Drag coeff. accuracy±2% vs wind tunnel
Mesh resolutiony⁺ < 1 (wall-resolved)
Aeroelastic toolMSC Nastran SOL 145
AIRFOIL PRESSURE DISTRIBUTION (NACA 0012) Cp=1 Cp DISTRIBUTION (AOA = 4°, Re = 3×10⁶) Upper Lower -Cp 0
Orbital Mechanics & Mission Design
Keplerian orbit determination, Hohmann transfer calculations, ΔV budget analysis, ground station coverage, eclipse fraction, and launch window optimisation. Software: GMAT, STK, custom Python propagators (J2 perturbation, atmospheric drag, solar pressure).
Orbit propagatorJ2 + J3 perturbation
ΔV accuracy±1 m/s
SoftwareGMAT / STK / Python
Orbit typesLEO, SSO, GEO, HEO, L1/L2
Vis-viva equationv² = GM(2/r − 1/a)
Hohmann ΔV LEO→GEO≈ 3.93 km/s total

Mission ΔV budget is the primary driver of propellant mass (Tsiolkovsky equation: Δv = Isp·g₀·ln(m₀/mf)). For a 500 kg GEO satellite with Isp=320s (bipropellant), LEO→GEO transfer requires ~830 kg of propellant — 62% of launch mass.

HOHMANN TRANSFER ORBIT Earth ΔV₁ ΔV₂ LEO ~500 km GEO 35786 km ΔV₁ = 2.46 km/s (LEO → transfer perigee) ΔV₂ = 1.47 km/s (transfer apogee → GEO) Total ΔV = 3.93 km/s ToF = 5.25 hours
Propulsion Systems Design
Chemical propulsion (monopropellant, bipropellant), electric propulsion (Hall thruster, gridded ion), and cold gas systems. Combustion chamber design (C* efficiency), nozzle design (Rao optimised bell), propellant budget, and pressurant system design.
BIPROPELLANT ROCKET ENGINE — RAO OPTIMISED BELL NOZZLE COMBUSTION CHAMBER LOX LH₂ ENGINE PERFORMANCE (LOX/LH₂) Isp (vac) = 450 s Thrust = 1500 kN Chamber P = 20 MPa Mixture ratio = 5.5 (O/F) Nozzle ratio ε = 77 (bell) C* efficiency = 98.5% Throat (A*) Bell nozzle (Rao) Ae/A*=77
450 s
Isp (LOX/LH₂)
3000 s
Isp (Hall Thruster)
20 MPa
Chamber Press.
98.5%
C* Efficiency
Thermal Protection System (TPS)
Re-entry vehicles experience peak heating rates of 1–10 MW/m² and stagnation temperatures of 5000–8000 K. TPS design involves ablative (PICA, SLA-561), TUFI tiles, and metallic TPS. Aerothermal analysis couples convective, radiative, and ablative heat transfer.
Peak heat rate10 MW/m² (stagnation)
Stagnation temp8000 K (equilibrium)
PICA density0.27 g/cm³
PICA use tempUp to 3600 K (ablative)
TUFI tile temp1600°C (reusable)
Analysis codeFIAT / CMA / custom
TPS LAYER STRUCTURE (PICA ABLATOR) HOT GAS (8000 K shock layer) PICA Ablator (25–50 mm) ρ = 0.27 g/cm³ · κ = 0.35 W/m·K Pyrolysis zone (char/virgin transition) Virgin PICA (unaffected core) SLA-561V filler / Bond coat CFRP Structural Shell (1.5 mm) 8000K 3000K 400K 300K RE-ENTRY HEATING TIMELINE Peak heating: t = 80s (interface alt.) Total heat load: 400 MJ/m² Recession rate: 0.3–0.8 mm/s Backwall T < 400 K throughout
Aerospace Composite Structures
CFRP (carbon fibre reinforced polymer) laminates designed using Classical Lamination Theory (CLT), progressive failure analysis, and damage tolerance assessments. Primary structures routinely achieve specific stiffness 5× aluminium at 40% the weight.
Fibre systemT700/T800/IM10 CF
MatrixRTM6 / Cytec 5215
E₁ (0° ply)165 GPa
E₂ (90° ply)9.5 GPa
G₁₂5.3 GPa
ν₁₂0.34
Fibre volume Vf60% (autoclave)
Areal density1600 g/m² (12-ply)
LAYUP SEQUENCE & CLT

[0/±45/90]₂s quasi-isotropic laminate — 16 plies, t = 2mm
CLT: [A B D] stiffness matrices from integration of Q̄k over thickness
Failure criteria: Hashin (fibre/matrix separation)
Progressive damage: ply degradation after 1st ply failure
Impact damage: BVID threshold 5–8 J (barely visible)
Residual strength: >1.5× DUL after BVID

Finite Element Analysis (FEA)
Full launch load FEA: static, dynamic (modal, random vibration, shock response spectrum), thermal, and buckling analyses. Models range from 10K to 10M DOF. Verification against analytical solutions (Roark's) and test data correlation within 5%.
FEA codeANSYS / MSC Nastran
Model size10K–10M DOF
Random vibrationPSD input (GEVS spec.)
Shock analysisSRS (Q=10, mil-std-1540)
Buckling analysisSOL 105 + knockdown
Test correlation<5% frequency error
FEA STRESS CONTOUR (SPACECRAFT BRACKET) σ_max 850 MPa 425 MPa 0 MPa BC: fixed F=50kN
Guidance, Navigation & Control (GN&C)
Attitude determination using star trackers, IMU sensor fusion (EKF), reaction wheel momentum management, and PID/LQR control law design. Orbit control using thruster pulse modulation and delta-V manoeuvres. AOCS design for 3-axis stabilised spacecraft.
Attitude accuracy±0.01° (3σ)
Attitude sensorStar tracker + IMU (EKF)
Actuators4 RWAs + 12 thrusters
Control lawLQR / PID hybrid
Slew rate2°/s (3-axis)
Settling time<60 s (large manoeuvre)
CONTROL SYSTEM BLOCK DIAGRAM
NAVIGATION Star/IMU EKF GUIDANCE Ref. trajectory CONTROL LQR/PID ACTUATORS RWA + RCS SPACECRAFT Dynamics Attitude feedback loop
Spacecraft Subsystem Engineering
End-to-end spacecraft design: mass budget, power budget (solar array + battery), link budget (downlink data rate), thermal control (MLI, heaters, heat pipes), and structural design from cubesat (1U) to 3-axis stabilised GEO platforms.
⚡ Power

Triple-junction GaAs solar array: 28% EOL efficiency. Li-ion battery (DOD 30%). Power budget from eclipse fraction and load profile. MPPT charge controller design.

📡 Comms

S/X-band link budget. EIRP, path loss (Friis), G/T, Eb/N₀ SNR margin. Data rates: 1 Mbps (S-band) to 100 Mbps (X-band). TT&C and mission data.

🌡 Thermal

MLI blanket design (εeff 0.02). Variable conductance heat pipes. Thermostat-controlled heaters. Radiator sizing: P_rad = εσAT⁴. Eclipse/sunlight cycling −40°C to +80°C.

🔩 Structure

Primary Al-7075 or CFRP isogrid panel structure. Random vibration test (GEVS 14.1 Grms). Shock test 3000g SRS. First mode freq >100 Hz.

🧭 ADCS

3-axis stabilised. Momentum wheel (0.1–2 N·m·s). Magnetorquer detumbling. Star tracker ±0.01° pointing. RWA desaturation via magnetic torquers.

💻 OBC

Radiation-hardened OBC (Leon3 FPGA). SpaceWire/CAN bus. RTOS (RTEMS/VxWorks). Fault detection, isolation, recovery (FDIR). SEU mitigation via TMR.

Aerospace Project Portfolio
Selected aerospace engineering projects across launch vehicles, satellites, re-entry vehicles, and atmospheric research platforms.
🚀 Launch Vehicle Bracket

Topology-optimised CFRP secondary structure bracket. 68% mass reduction vs baseline Al casting. FEA certified to 200% DUL. First mode: 250 Hz.

🛰 6U CubeSat Structure

Al-7075 monolithic 6U cubesat structure for LEO Earth observation. Mass: 280g. Passed qualification testing (14.1 Grms random, 3000g SRS). Delivered 2024.

🔥 Re-entry Capsule TPS

PICA ablator TPS sizing for 11 km/s entry velocity ballistic capsule. Peak heat rate 8.5 MW/m². FIAT analysis with 2.5× mass margin. 40mm ablator thickness.

✈ Hypersonic Waverider

Mach 6 waverider aeroshell CFD analysis. L/D = 3.8 at design point. Scramjet intake compression ramp optimisation. Titanium-UHTC leading edge TPS design.

🌍 GEO Platform AOCS

3-axis AOCS design for 5000 kg GEO telecom satellite. LQR attitude control, 4-RWA momentum management, and dual xenon propulsion ΔV for 15-year station-keeping.

🛸 Airborne Radar Dome

CFRP sandwich radome structure. RF transparency >97% at X-band. Compressive strength 4500 kPa. EUROCAE ED-14G environmental qualification. ±60°C operational.

Mission-Critical Aerospace Engineering

From preliminary design review (PDR) to launch readiness review (LRR), Neo Materials provides full mission lifecycle aerospace engineering support.

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