Space Systems Engineering

Space systems engineering encompasses the design, development, and operation of spacecraft and satellites. Unlike terrestrial systems, spacecraft must operate autonomously in the harsh environment of space for extended periods, often without opportunity for maintenance. This requires careful integration of multiple subsystems to ensure mission success.

Satellite Subsystem Architecture

Primary Subsystems

A typical satellite contains the following subsystems:

1. Structure: Provides mechanical support and protection
2. Power: Solar arrays, batteries, power distribution
3. Thermal: Temperature control and heat management
4. Propulsion: Orbit raising, station-keeping, attitude control
5. GNC (Guidance, Navigation, Control): Attitude determination and control
6. Communication: Telemetry, tracking, and command (TT&C)
7. Payload: Mission-specific instruments and sensors
8. Data Handling: On-board computers and data management
9. Mechanisms: Solar array drive, antenna pointing, deployment devices

System Integration Challenges

Spacecraft must meet stringent requirements:

• Mass budget: Every gram is important
• Power budget: Limited by solar panel area and efficiency
• Thermal budget: Must operate in extreme temperature variations
• Reliability: Often redundant systems to ensure mission success
• Environmental: Survive launch loads and space environment

Structural Design

Materials Selection

Spacecraft structures use lightweight, high-strength materials:

• Aluminum alloys: 6061-T6, 7075-T6 for general structure
• Carbon fiber composites: High stiffness-to-weight ratio
• Titanium: High strength, good thermal properties
• Invar: Low coefficient of thermal expansion for precision components

Structural Configurations

• Box structures: Common platform configuration
• Truss structures: High stiffness for large apertures
• Deployable structures: Solar arrays, antenna reflectors

Loads Analysis

Spacecraft experience loads during:

• Launch: High acceleration and vibration
• Acoustic environment: Severe sound pressure levels
• Thermal cycling: Extreme temperature changes
• Micrometeoroid and debris: Ballistic impact protection

Power Systems

Solar Array Power Generation

Solar power in space follows the relationship:

Where:

• $A$ = solar panel area
• $I_0$ = solar constant (1367 W/m² at 1 AU)
• $\eta$ = panel efficiency
• $\theta$ = angle between normal vector and sun direction

For Earth orbits, the solar constant varies slightly based on solar activity and distance from sun.

Battery Technology

• Nickel-Hydrogen (NiH₂): Mature technology, 1000+ cycles
• Lithium-Ion (Li-ion): Higher energy density, lighter weight
• Lithium-Polymer (Li-Po): Flexible form factor, newer technology

Battery state of charge calculation:

Where $V_{oc}$ is open-circuit voltage, and $V_{max}$, $V_{min}$ are fully charged and discharged voltages.

Power Management

The power system must handle:

• Peak power demands: During high-activity operations
• Average power: Sustained power requirements
• Energy storage: Battery capacity for eclipse periods
• Power conditioning: Voltage regulation and filtering

Thermal Control Systems

Heat Transfer in Space

In space, heat transfer occurs primarily through:

• Radiation: $Q = \varepsilon \sigma A (T_1^4 - T_2^4)$
• Conduction: Through structural elements
• Convection: Negligible in vacuum

Where $\varepsilon$ is emissivity, $\sigma$ is Stefan-Boltzmann constant, and $T$ is temperature in Kelvin.

Passive Thermal Control

• Multi-layer insulation (MLI): Reflects thermal radiation
• Optical surface reflectors: Control solar absorption
• Heat pipes: Passive heat transfer devices
• Phase change materials: Thermal energy storage

Active Thermal Control

• Heaters: Maintain component temperatures during cold periods
• Thermal louvers: Variable emissivity systems
• Loop heat pipes: Active heat transport
• Thermoelectric coolers: Peltier effect cooling

Thermal Analysis

Steady-state thermal balance:

Where $Q_{generated}$ includes power dissipation from electronics and solar absorption.

Propulsion Systems

Chemical Propulsion

• Monopropellant: Hydrazine (N₂H₄), simple, reliable
• Bipropellant: MMH/NTO, higher performance
• Solid propellant: Pre-packed fuel, simple but not throttleable

Specific impulse for chemical rockets:

Where $T_c$ and $P_c$ are chamber temperature and pressure.

Electric Propulsion

• Ion thrusters: High efficiency, low thrust
• Hall effect thrusters: Moderate efficiency, higher thrust than ion
• Resistojet: Simple electric thrusters with moderate efficiency

Propulsion Selection Factors

• Mission duration: Electric for long-duration missions
• Delta-v requirements: Chemical for high delta-v needs
• Power availability: Electric requires significant power
• Thrust level: Chemical for rapid maneuvers

Guidance, Navigation and Control (GNC)

Attitude Determination

Spacecraft use multiple sensors to determine orientation:

Star Trackers

Most accurate attitude sensors, identify stars in field of view to determine orientation.

Sun Sensors

Simple, reliable sensors used for coarse pointing and safe mode operation.

Earth Horizon Sensors

Detect Earth's limb for nadir-pointing missions.

Gyroscopes

Measure angular rates for attitude propagation between measurements.

Attitude Control

Control Methods

• Momentum wheels: Precise pointing, continuous control
• Control moment gyros (CMG): High authority for large spacecraft
• Magnetic torquers: Use Earth's magnetic field, limited applications
• Thrusters: High control authority, fuel consumption

Control Laws

Three-axis attitude control using quaternions:

Where $\Omega(\omega)$ is the skew-symmetric matrix form of angular velocity vector $\omega$.

Proportional-Derivative control:

Where $\sigma_{br}$ is the modified Rodrigues parameters error and $\omega_{br}$ is the body rate error.

Communications Systems

Antenna Types

• Patch antennas: Low profile, circular polarization
• Helical antennas: Good circular polarization, wide bandwidth
• Parabolic reflectors: High gain for deep space missions
• Phased arrays: Electronic beam steering capability

Link Budget Analysis

Where:

• $P_r$ = received power
• $P_t$ = transmitted power
• $G_t$, $G_r$ = transmitter and receiver gains
• $L_{fs}$ = free space loss
• $L_{other}$ = other losses

Free space loss:

Where $d$ is distance and $\lambda$ is wavelength.

Communication Protocols

• TM/TC: Telemetry and telecommand protocols
• CCSDS: Consultative Committee for Space Data Systems standards
• SpaceWire: High-speed data communication standard
• CFDP: CCSDS File Delivery Protocol for data file transfer

Data Handling Systems

Flight Computer Architecture

Modern spacecraft use:

• Redundant processors: Triple modular redundancy (TMR) or dual-string
• Radiation-hardened components: To survive space radiation
• Real-time operating systems: Deterministic task scheduling
• Fault detection/isolation/recovery (FDIR): Autonomous anomaly handling

Memory Systems

• SRAM: Fast, radiation tolerant, expensive
• EEPROM: Non-volatile, limited write cycles
• Flash: Higher density, sensitive to radiation
• FRAM: Ferroelectric RAM, radiation tolerant, unlimited writes

Data Storage

Data compression is crucial for limited downlink opportunities.

Ground Systems

Ground Station Network

• Primary stations: Command and control
• Backup stations: Redundancy for continuous coverage
• Science stations: Data reception from instruments
• Tracking stations: Orbit determination and monitoring

Mission Operations

• Mission Control Center: Real-time operations
• Payload Operations Center: Scientific data processing
• Network Operations Center: Communication system management

Orbit Determination

Precise orbit determination uses:

Where $\mathbf{a}_{pert}$ includes perturbations from Earth's gravity field, lunar/solar attraction, solar radiation pressure, and atmospheric drag.

Real-World Application: CubeSat Power System Design

Consider the power system design for a 3U CubeSat mission for Earth observation.

Power Budget Analysis

import math

# CubeSat parameters
satellite_size = 3  # U (1U = 10x10x10 cm)
total_volume = 3000  # cm³ (3U CubeSat)
available_volume_solar = 200  # cm² (top and side panel surface area available for solar cells)

# Solar cell parameters (typical for CubeSats)
solar_cell_efficiency = 0.28  # 28% efficiency for advanced GaAs cells
solar_constant = 1367  # W/m² in space
solar_incidence_factor = 0.85  # Account for orbital geometry and panel orientation

# Power requirements (typical CubeSat mission)
power_bus_voltage = 7.5  # V (common CubeSat power bus)
duty_cycle_sunlight = 0.57  # 57% of time in sunlight, 43% in eclipse
eclipse_duration = (24 * 3600) * (1 - 0.57)  # seconds in eclipse

# Calculate available solar power
available_area = available_volume_solar / 10000  # Convert to m²
available_power = (available_area * solar_constant * solar_cell_efficiency * 
                   solar_incidence_factor)  # Units: Watts

print(f"Available solar panel area: {available_area*10000:.1f} cm²")
print(f"Available solar power: {available_power:.2f} W")
print(f"Time in eclipse: {eclipse_duration/3600:.2f} hours per orbit")

# Calculate battery requirements
# Assume peak power consumption during eclipse
peak_consumption = available_power * 0.7  # Use 70% of available power
eclipse_energy_required = peak_consumption * eclipse_duration  # Watt-seconds

# Battery capacity needed (with margin)
battery_margin = 1.5  # 50% margin for degradation and temperature effects
required_battery_capacity = eclipse_energy_required * battery_margin / 3600  # Convert to Wh

print(f"Peak power consumption: {peak_consumption:.2f} W")
print(f"Energy required per eclipse: {eclipse_energy_required/3600:.2f} Wh")
print(f"Required battery capacity: {required_battery_capacity:.2f} Wh")

Thermal Considerations

Spacecraft thermal design must account for the significant temperature variations in space.

Orbital Mechanics Impact

Orbital parameters significantly affect both power and thermal subsystems.

Your Challenge: Satellite Thermal Control Design

Design a thermal control system for a satellite in sun-synchronous orbit that maintains all components within operational temperature ranges.

Goal: Calculate the required thermal control system parameters to maintain proper temperatures for a satellite's critical components.

Satellite Parameters

# Satellite specifications
satellite_area_total = 4.0  # m² (total surface area)
satellite_mass = 250  # kg
satellite_surface_absorptivity = 0.6  # Solar absorptivity
satellite_surface_emissivity = 0.8  # Thermal emissivity

# Orbital parameters (Sun-synchronous orbit)
altitude = 800000  # m (800 km altitude)
beta_angle = 45  # degrees (angle between orbit plane and sun direction)
solar_input_factor = math.cos(math.radians(90 - beta_angle))  # Direct solar input factor

# Component specifications
electronics_power = 200  # W (total internal heat generation)
thermal_capacity = 1200  # J/(kg·K) (satellite average thermal capacity)
required_temperature_range = (-5, +40)  # Celsius (operational range)

# Thermal control parameters
heater_efficiency = 0.9  # Heater power effectiveness factor
heat_pipe_capacity = 50  # W (heat transport capability)
mlc_emissivity = 0.7  # Multi-layer insulation emissivity

# Space environment
solar_constant = 1367  # W/m²
boltzmann_constant = 5.67e-8  # W/(m²·K⁴)

Design a thermal control system that can maintain the satellite within its operational temperature range throughout the orbit.

Hint:

• Calculate the thermal balance equation: internal heat generation + absorbed solar radiation = radiated heat + heat stored
• The radiated heat is given by: εσA(T⁴ - T_space⁴) where T_space ≈ 4K
• For steady state: Heat in = Heat out
• Calculate the required heater power and passive thermal control measures

# TODO: Calculate thermal balance parameters
absorbed_solar_power = 0  # W (solar radiation absorbed by satellite)
radiated_heat_power = 0   # W (heat radiated to space)
heater_power_required = 0 # W (power for heaters to maintain minimum temperature)
steady_state_temperature = 0  # Celsius (equilibrium temperature without heaters)

# Calculate temperature range during orbit
max_temperature = 0  # Celsius
min_temperature = 0  # Celsius

# Print results
print(f"Absorbed solar power: {absorbed_solar_power:.2f} W")
print(f"Steady state temperature (no heaters): {steady_state_temperature:.1f} °C")
print(f"Required heater power: {heater_power_required:.2f} W")
print(f"Temperature range: {min_temperature:.1f} to {max_temperature:.1f} °C")
print(f"Operational temperature range: {required_temperature_range[0]} to {required_temperature_range[1]} °C")

# Check if design meets requirements
design_meets_requirements = (min_temperature >= required_temperature_range[0] and 
                            max_temperature <= required_temperature_range[1])
print(f"Design meets temperature requirements: {design_meets_requirements}")

What additional thermal control measures would you implement to ensure the satellite operates within safe temperature limits under all orbital conditions?