Week 7 Intro To UAS 270, Comprehensive Study Guide on UAS Powerplants, Electrical Architectures, and Motor Control

Photovoltaic Cells and Operating Principles

• Definition and Fundamental Function:

  • Solar cells are solid-state semiconductor devices designed to convert electromagnetic radiation (light energy) directly into usable electrical energy through the photovoltaic effect.

• Semiconductor Architecture and Charge Carriers:

  • N-type Semiconductor (Silicon): Doped to host an excess of free electrons (negatively charged carriers).

  • P-type Semiconductor (Silicon): Doped to host an excess of electron "holes," which represent the absence of negative charges and behave as positively charged carriers.

  • The p-n Interface: The boundary between the p-type and n-type silicon layers that establishes an internal electrostatic field.

• Photovoltaic Conversion Mechanism:

  • Photons from incoming sunlight strike the solar cell and are focused primarily on the n-type semiconductor layer.

  • Absorbed photon energy excites bound electrons, providing the energy required to "free up" electrons by driving them across the p-n junction from the p-type semiconductor to the n-type semiconductor.

  • The accumulation of energized free electrons in the n-type silicon establishes an electric potential.

  • When an external circuit connects the two layers, these free electrons migrate out through the circuit to deliver power to an electrical load, ultimately returning to the p-type semiconductor to recombine with positive holes.

• Electrical Output Characteristics:

  • A single standard silicon solar cell typically generates an open-circuit potential of approximately $0.58\,\text{V}$.

  • The typical current density generated by a single cell is approximately $35\,\text{mA/cm}^2$.

• Circuit Configurations for Array Scaling:

  • Series Connection: Connects individual cells end-to-end to scale and increase the total output voltage ($V_{\text{total}} = V_1 + V_2 + \dots + V_n$).

  • Parallel Connection: Connects cells side-by-side to scale and increase the total output current ($I_{\text{total}} = I_1 + I_2 + \dots + I_n$).

  • Series-Parallel Combinations: Arrays wired in hybrid networks to simultaneously meet distinct voltage and current operational requirements.

Solar Integration in Unmanned Aerial Systems (UAS)

• Airframe Compatibility and Surface Area Limitations:

  • Generating sufficient operational current and voltage to power an entire propulsion and avionics system demands an extensive surface collector area.

  • Consequently, solar panels are almost exclusively integrated into fixed-wing aircraft architectures, where wings and upper empennage surfaces provide the requisite planar area.

  • Multirotor airframes generally lack adequate surface area to host sufficient solar arrays relative to their power consumption.

• Power-to-Area Scaling:

  • The total collector area required scales directly with the electrical power demanded by the aircraft systems.

  • Design Example: Delivering continuous power of $2000\,\text{W}$ under direct, uninterrupted sunlight demands approximately $200\,\text{ft}^2$ of active solar panel collector area.

• Conversion Efficiency and Structural Sizing:

  • Commercial-grade silicon solar panels feature an efficiency rating of only approximately $21\%$, meaning only about one-fifth of incident radiant light is successfully transduced into electricity.

  • Because of this low conversion efficiency, solar-powered UAS must be constructed with exceptionally large wingspans and high-aspect-ratio wings to maintain adequate power-to-area and power-to-mass ratios.

• Weight and Redundancy Considerations:

  • Aircraft weight budgets must accommodate the parasitic mass of the solar panels, inter-cell cabling, structural substrates, and power regulation electronics.

  • Solar panels cannot deliver uniform continuous power under all flight conditions; therefore, rechargeable onboard batteries are essential to serve as an energy buffer and redundant failsafe during banking, cloud cover, or nighttime flight.

• Incidence Angle and Solar Alignment:

  • Peak power generation occurs exclusively when solar panels are oriented in direct sunlight with normal incidence ($90^\circ$ angle, where light rays are perpendicular to the panel face).

  • Any angular deviation away from normal incidence (due to aircraft attitude, yaw, roll, or low sun elevation) decreases the projected surface area intercepting the solar flux, significantly attenuating output power.

Solar Insolation and Power Sizing Calculations

• Solar Insolation Definition:

  • Solar insolation is defined as the measure of total incoming solar radiant energy impinging upon a specific surface area over a given period.

  • The baseline standard for maximum solar insolation at the Earth\'s surface at solar noon under clear, cloudless sky conditions is approximately $1000\,\text{W/m}^2$.

• Environmental and Geographic Variables Affecting Insolation:

  • The axial tilt and orbital position of the Earth.

  • Diurnal rotation (time of day).

  • Geographic latitude.

  • Atmospheric conditions, including cloud density, relative humidity, particulate pollution, and suspended aerosols.

• Sizing Calculation: Replacing a Lithium-Polymer Battery with Solar Power:

  • Scenario: Calculate the minimum active solar cell area required to continuously replace a 3-cell ($\text{3S}$) Lithium-Polymer (LiPo) battery operating at $12.6\,\text{V}$ and discharging $5\,\text{A}$.

  • Assumptions: Peak solar insolation ($I_{\text{solar}} = 1000\,\text{W/m}^2$) and commercial silicon solar cell efficiency ($\eta = 21\% = 0.21$).

  • Step 1: Compute Required Electrical Power ($P$):     $P = I \times V$     $P = (5\,\text{A}) \times (12.6\,\text{V}) = 63\,\text{W}$

  • Step 2: Relate Power to Insolation, Panel Area, and Efficiency:     $P = I_{\text{solar}} \times \text{Area} \times \eta$

  • Step 3: Solve for Required Collector Area ($\text{Area}$):     $\text{Area} = \frac{P}{I_{\text{solar}} \times \eta}$     $\text{Area} = \frac{63\,\text{W}}{(1000\,\text{W/m}^2) \times (0.21)} = \frac{63\,\text{W}}{210\,\text{W/m}^2} = 0.3\,\text{m}^2$

  • Step 4: Convert to Imperial Units:     $0.3\,\text{m}^2 \approx 3.2\,\text{ft}^2$

  • Conclusion: Supplying $63\,\text{W}$ to replace the discharge of a $\text{3S}$ LiPo battery requires a minimum solar panel surface area of $0.3\,\text{m}^2$ ($3.2\,\text{ft}^2$) operating at peak normal incidence.

Hydrogen Fuel Cell Technology

• Electrochemical Operating Architecture:

  • A hydrogen fuel cell operates analogously to an open electrochemical battery, continuously generating electricity via chemical reactions as long as reactant gases are supplied.

  • Essential assembly layers include an anode (negative electrode), a cathode (positive electrode), and a central Polymer Electrolyte Membrane (PEM).

• Step-by-Step Chemical Conversion Process:

  • 1. Fuel Delivery: Hydrogen gas ($H_2$) is directed across field flow plates to the anode on one side of the fuel cell; simultaneously, an oxidant ($O_2$ from atmospheric air or pure oxygen) is channeled through flow plates to the cathode on the opposite side.

  • 2. Anodic Catalytic Splitting: At the anode backing layer, a platinum catalyst facilitates the oxidation of hydrogen, splitting each molecular hydrogen molecule into positively charged hydrogen ions (protons, $H^+$) and free electrons ($e^-$):     $H_2 \rightarrow 2H^+ + 2e^-$

  • 3. Charge Separation and Current Generation: The Polymer Electrolyte Membrane (PEM) acts as an ion-selective barrier; it permits only positively charged protons ($H^+$) to pass directly through the membrane to the cathode. Negatively charged electrons ($e^-$) are blocked by the membrane and forced to travel along an external electrical circuit, generating the useful DC electric current that powers onboard avionics and motors.

  • 4. Cathodic Reduction and Water Synthesis: At the cathode, arriving electrons from the external circuit and migrating $H^+$ protons combine with oxygen gas ($O_2$). Two $H^+$ ions bond with one reduced oxygen atom to synthesize pure water ($H_2O$):     $\frac{1}{2}O_2 + 2H^+ + 2e^- \rightarrow H_2O$

  • Byproducts: The primary exhaust/waste byproduct is clean water ($H_2O$) and negligible residual heat, making it an entirely zero-emission local powerplant.

• Electrical Scaling and Output Parameters:

  • Individual Cell Voltage: A single PEM fuel cell generates a nominal output voltage between $0.5\,\text{V}$ and $0.8\,\text{V}$.

  • Stacking Configuration: To achieve operational voltages required for aerial propulsion, individual fuel cells are connected in series into a "fuel cell stack" (e.g., a stack containing 10 series cells produces a cumulative potential between $5\,\text{V}$ and $8\,\text{V}$).

  • Current Capacity Regulation: Total current output ($\text{Amperes}$) is directly governed by the active cross-sectional surface area of the membrane plates; a larger plate area provides more catalytic sites to oxidize hydrogen and liberate a greater volume of electrons per second.

  • Interactive Educational Reference: Lawrence Hall of Science SEPUP Fuel Cell Simulation: https://sepup.lawrencehallofscience.org/sepup-fuel-cell-simulation/.

Electric Motor Control and Electronic Speed Controllers (ESCs)

• Purpose of the Electronic Speed Controller:

  • Brushless DC (BLDC) motors cannot self-regulate their speed or phase sequencing when supplied with plain direct current.

  • An Electronic Speed Controller (ESC) provides precise speed and torque regulation of the brushless motor shaft (and thereby propeller thrust) by converting DC battery power into a rapid sequence of pulsed voltage waveforms across the motor phases.

• Pulse Width Modulation (PWM) and Duty Cycle Principles:

  • Rather than stepping down voltage linearly, the ESC switches full voltage rapidly on and off via high-frequency Pulse Width Modulation (PWM).

  • Duty Cycle: The ratio of active conduction time ("on" duration) relative to the total switching period (cycle period):     $\text{Duty Cycle} = \frac{T_{\text{on}}}{T_{\text{on}} + T_{\text{off}}} \times 100\%$

  • Modulation Levels:

  • $25\%$ Duty Cycle: Voltage is pulsed "on" for $25\%$ of the time and remains "off" for $75\%$; produces low average voltage, delivering low power and slow propeller rotation.

  • $50\%$ Duty Cycle: Voltage is pulsed "on" for $50\%$ of the time and "off" for $50\%$, supplying medium power.

  • $75\%$ Duty Cycle: Voltage is pulsed "on" for $75\%$ of the period and "off" for $25\%$, delivering high power and rapid propeller rotation.

  • $100\%$ Duty Cycle: Continuous "on" state, supplying peak possible power and maximum motor RPM.

• Thermal Losses and Cooling Requirements:

  • The internal switching MOSFETs inside an ESC generate the highest levels of waste heat when actively transitioning states (switching regimes between $0\%$ and $100\%$ duty cycle).

  • Because drone power systems continuously operate at partial throttle intermediate duty cycles, the ESC generates significant heat.

  • Adequate passive heat sinking or forced convective airflow is required to prevent thermal throttling or component failure.

• Internal Circuit Anatomy of an ESC:

  • Input Capacitor: Stabilizes the main DC power bus against severe voltage spikes caused by motor switching transients.

  • Linear BEC (Battery Eliminator Circuit): Voltage regulator circuit providing a stepped-down $5\,\text{V}$ supply to power receivers and servos without needing a secondary battery.

  • ATmega8 MCU: Dedicated central microcontroller that runs the ESC firmware, calculates timing logic, and commands the switching stages.

  • Linear MCU Regulator: Delivers clean, ripple-free voltage specifically to power the onboard ATmega8 processor.

  • External Oscillator: Precision clock source (e.g., $16.0\,\text{MHz}$) that maintains cycle-accurate timing for PWM wave generation and phase commutation.

  • Discrete FET Driving Components: Gate driver transistors and switching circuitry that drive the high-power Field Effect Transistors (FETs).

  • Phase Sense Voltage Dividers & Sense Filtering Capacitors: Resistive and capacitive networks that tap into motor phase leads to measure Back-Electromotive Force (Back-EMF), enabling sensorless rotor position detection.

  • Pads for ISP Programming: In-System Programming pin terminals used to flash or update firmware on the microcontroller.

  • Wiring Interfaces: Main DC battery input leads (positive and ground), 3-wire signal lead (PWM signal input, BEC $5\,\text{V}$ output, ground), and 3-phase motor output leads.

Fixed-Wing UAS Electronics Architecture

• Baseline System Architecture:

  • A fixed-wing unmanned aircraft requires a foundational set of electromechanical and radio-frequency components to achieve controlled flight.

• Primary Components and Respective Functions:

  • Motor: Brushless electric motor connected to a propeller to provide forward thrust for aerodynamic lift.

  • Electronic Speed Controller (ESC): Receives throttle commands from the radio receiver and drives the motor with modulated high-current 3-phase power.

  • Battery: High-discharge chemical energy source (such as Lithium-Polymer) that supplies DC electric power to the entire aircraft.

  • Receiver (Rx): Radio frequency module that receives telemetry and flight control commands transmitted from the handheld Remote Control or Ground Control Station (GCS).

  • Servos: High-torque miniature electromechanical rotary actuators wired directly to receiver signal channels; they physically displace aerodynamic control surfaces:

  • Ailerons: Modulate asymmetric wing lift to control roll.

  • Elevator: Modulates horizontal tail incidence to control pitch.

  • Rudder: Modulates vertical tail incidence to control yaw.

• Power and Signal Routing:

  • The battery connects directly to the ESC via high-gauge DC wiring.

  • The ESC's integrated Battery Eliminator Circuit (BEC) outputs stepped-down $5\,\text{V}$ power over a 3-pin lead to energize the receiver's internal power bus.

  • The receiver distributes power and sends individual PWM positioning pulses to the control surface servos while routing throttle commands back to the ESC.

Multirotor (Quadcopter) Electronics Architecture

• Baseline System Architecture:

  • Multirotor aircraft rely on differential thrust across multiple motors rather than aerodynamic control surfaces to achieve attitude stabilization, heading control, and translational motion.

• Primary Components and Functional Relationships:

  • Motors (4x): Four brushless DC outrunner motors (e.g., $1400\,\text{KV}$ rating) driving matched fixed-pitch propellers to generate vertical thrust and torque control.

  • Electronic Speed Controllers (4x ESCs): Four discrete ESC modules wired individually to each motor, enabling independent, high-frequency RPM adjustments.

  • Battery: Lithium-Polymer (Li-Po) battery pack equipped with an XT60 high-current connector, routing power across a central parallel harness to all four ESCs.

  • Flight Controller (e.g., ArduPilot Mega 2.8 Board): The central computational autopilot computer; contains internal inertial measurement units (gyroscopes and accelerometers). It computes aircraft state hundreds of times per second and dispatches individual PWM speed commands to the four ESCs to stabilize and maneuver the airframe.

  • Remote Control Receiver: Intercepts pilot navigation signals over a $2.4\,\text{GHz}$ antenna array and forwards these channel inputs to the flight controller.

  • GPS Module: External satellite receiver interfaced with the flight controller to provide absolute latitude, longitude, and altitude telemetry for autonomous waypoint following, geofencing, and return-to-launch (RTL) capabilities.

Non-Electric and Hybrid Powerplants for UAS

• Internal Combustion and Turbine Power Systems:

  • For long-range and heavy-payload small UAS (sUAS), propulsion may be delivered by small internal combustion engines (ICE) or micro-turbine/hybrid turbojet engines rather than electric motors.

  • Regardless of primary propulsion type, all UAS still require clean electrical energy to power avionics, autopilots, radio links, and sensor payloads.

• Engine Components of a Gas-Powered Unit:

  • Power Generator Mounting Brackets: Secure the engine and generator assembly to the aircraft chassis.

  • Engine Crankcase with Integrated Control Device: Houses the crankshaft and engine management electronics for ignition and throttle control.

  • Cylinder of the Internal Combustion Engine: The thermal combustion chamber where fuel-air mixtures are ignited.

  • Spark Plug: Delivers timed high-voltage electrical ignition to the cylinder.

  • Electric Power Unit: Generator module mechanically linked to the engine shaft to produce electrical current.

• Liquid Fuel Energy Density Advantage:

  • Liquid hydrocarbon fuels (such as gasoline, diesel, and aviation kerosene) possess an energy density up to $50\times$ higher than modern chemical lithium-based batteries.

  • This energy density differential enables combustion-powered UAS to achieve substantially longer endurance, higher range, and lower overall airframe weights than battery-dependent equivalents.

• Onboard Generator Principles:

  • Rather than carrying large non-rechargeable battery arrays, larger combustion UAS utilize an onboard electrical generator driven by the engine.

  • The engine provides mechanical rotational power above and beyond the baseline requirements for thrust/propulsion, which drives the generator.

  • Operating Mechanism: An electric generator is fundamentally an electric motor operated in reverse: an external mechanical force (the engine) rotates the central shaft, spinning magnets past copper windings to induce a voltage and electrical current via Faraday's law.

  • Generator units are purpose-built and wound to maximize electrical generation efficiency rather than mechanical torque output.

• Hybrid Energy Architectures:

  • Many combustion UAS employ a hybrid configuration consisting of an onboard engine generator coupled with a secondary rechargeable battery bank.

  • The generator powers avionics and charges the battery bank during cruising flight.

  • The battery serves as a dynamic buffer during transient high-demand power draws and provides uninterruptible emergency power to sustain flight control systems in the event of an engine failure or flameout.

  • Multimedia Reference: Hybrid Turbojet Generator demonstration video: https://www.youtube.com/watch?v=9xw3neYuehw.