Browse Topic: Flight dynamics
The Sikorsky S-92® helicopter fleet, representing more than 300 aircraft and 2.6 million flight hours, is relied upon to support a large range of important missions across the globe. In previous efforts, a high-fidelity CFD-CSD based full-aircraft simulation methodology, co-simulated with production FCS, was developed and applied to model both coaxial aircraft and single main/tail rotor configurations (Refs. 1-5). The CFD solver is based on the CREATE™-AV HELIOS toolset (Ref. 6) and the CSD solver is based on Rotorcraft Comprehensive Analysis System (RCAS) (Ref. 7). The current paper further correlated the CoSim methodology (Ref. 1) with the S-92® helicopter flight-test database at both hover, cruise and edge-of-envelope maneuver flight conditions. The consistent correlations for flight dynamics, static and fatigue component loads at conditions across the flight envelope demonstrate the reliable predictive capability of the high-fidelity CoSim methodology to be-used as a virtual digital flight test and to support advanced design at early stage.
The certification of highly integrated electric Vertical Take-Off and Landing (eVTOL) aircraft requires a rigorous bridge between simulation and flight reality. This paper presents the Joby Disturbance Generator, a high-integrity software framework natively integrated into the aircraft flight control system. The system utilizes a deterministic state machine to inject a library of signals, ranging from standard doublets and chirps to complex waveforms, directly into internal control loops. Applications include frequency sweeps for stability margin extraction and structural mode identification, time-domain inputs for handling qualities assessment, synthetic fault injection for redundancy management verification, and precise loads model validation. The system continuously monitors vehicle health, automatically aborting test points upon detecting genuine failures. For loads validation, it coordinates temporary relaxation of flight envelope protections with precise disturbance injection. This methodology accelerates development by shifting risk from flight execution to software verification, providing deterministic, data-driven evidence for certification. Flight test results demonstrating these capabilities are presented.
Fault detection in autonomous VTOL aircraft is critical because even minor degradations can quickly destabilize multirotor vehicles in safety-critical environments. However, real-flight fault detection remains challenging due to sensor noise, environmental disturbances, and the nonlinear aeromechanics of multirotor platforms. This study proposes a comprehensive machine-learning framework for rotor fault detection, isolation, and severity prediction using real flight data. A convolutional neural network (CNN) architecture is developed to learn spatio-temporal patterns from multivariate flight dynamics, enabling direct inference of both the faulty rotor and its damage level. The framework is first validated using simulated data generated by our in-house flight dynamic model. Next, to verify the framework using real flight data, a hexcopter was designed, fabricated and flight tested for both nominal and faulty cases by introducing controlled blade-tip breakage. The trained model achieves rotor-wise fault classification accuracies above 99% and sample-wise severity estimation accuracy of 96% within a ±1% tolerance in experimental data, demonstrating strong generalization and supporting real-time health monitoring for autonomous VTOL systems.
NASA's successful demonstration of powered flight on Mars through the Ingenuity Helicopter, as part of the Mars 2020 Perseverance rover mission, has led to the development of next generation Martian rotorcraft. The future of Martian rotorcraft has evolved to include high payload-carrying vehicles to possibly contribute to planetary science missions, which will require improved flight dynamics and rotor aerodynamic performance to fly at nominally high forward flight speeds and at higher flight altitudes. To ensure the feasibility and viability of successful mission performance, it is also critical to mature the structural design for advanced Martian rotorcraft to bridge the gap between the best practices of the spacecraft and aircraft communities. This paper focuses on the structural analysis of a Mars Science Helicopter (MSH) blade using finite element methods. Multiple loading conditions including launch and operational flight were applied to investigate the blade’s structural integrity. The blade’s modal natural frequencies were also analyzed to investigate the blade's dynamic behavior.
The paper discusses the design and high-fidelity flight dynamics modeling of a 13-lb lift-plus-cruise unmanned aerial vehicle (UAV) using Rotorcraft Comprehensive Analysis System (RCAS) in order to (1) better understand its physics of flight during a wide range of maneuvers, and (2) provide insight into the fidelity needed to achieve quantitative accuracy when compared to flight test data. Wind tunnel tests of the full aircraft were performed at a 65% scale to provide lookup tables for the flight dynamics model. Flight test data was collected while providing high control inputs to excite a variety of dynamic states in hovering and cruising modes to systematically validate the physics model. Near quantitative agreement was observed between the model predictions and test data during hover; however, the predictions began to disagree at higher forward cruising speeds. To address the discrepancy between the prediction and experiment, the flight dynamics model was improved by learning a correction from flight test data using a neural network. This hybrid physics plus data-driven approach reduced the error between the physics model and experiment by 74% and only needing 12 minutes of flight data for training. This hybrid methodology presents an alternate approach to high fidelity modeling which only needs a relatively small amount of flight test data.
Previous researchers developed equations to model the induced flow on a 2D airfoil in the finite-state as opposed to the closed-form. Those models, however, were limited in that they could not handle an oscillating free stream that became negative. Recently, a new model was developed to include a single factor to carry the effects of the free stream changing signs. In developing this model, a Floquet instability was discovered at the instant when the flow changes direction. The effect of the instability grows with increasing number of oscillations of the sign of the free stream. The effects can be limited depending on the parameters of the flow. In this paper, the previous 2D model is amended to include a term that considers the effects of the induced flow from all previous vorticity segments that have been generated from each oscillation of the flow. This paper details the beginnings of the testing on the stability limits of the theory, based on changing the parameters of the free stream, airfoil, and timing. It is the intention of this research to further investigate the limits of this model in reversing flow in hopes of using the lessons learned to extend 3D finite-state models that are currently incapable of handling cases where the sign of the free-stream velocity changes as a result of the rotor reentering its own wake such as when a helicopter or quadcopter quickly descends after ascending to avoid an obstacle.
This paper presents the design, development, and subscale flight testing of an optionally-autonomous lift-plus-cruise (LPC) eVTOL aircraft for emergency response missions that bridges the gap between existing aerial capabilities and the needs of first responders. A 4+1 LPC configuration consisting of four vertical lift propellers and a single pusher propeller was selected to balance hover performance and cruise efficiency. The vehicle is sized around a 600 lbs gross takeoff weight with a 125 lbs payload capacity. VTOL and Pusher propeller blades were optimized using parametric studies, resulting in a high Figure of Merit and propulsive efficiency. Trim analysis demonstrates efficient hover to cruise transition, lift-to-drag ratios of 10-11 between 70-90 knots, and propulsive efficiency exceeding 0.9 at the cruise speed of 100 knots. The subscale configuration utilized a simulation framework for trim and optimization of flight control laws, which were subsequently implemented on a 1/3-scale subscale demonstrator. Subscale flight tests showed stable hover and robust trajectory tracking under wind disturbances throughout the flight envelope, which demonstrates the feasibility of the proposed LPC architecture.
Characterization of rotor–rotor wake interactions and their influence on flight dynamics is an important step toward advancing control system design and evaluating the performance of next-generation Mars multirotors. In this work, a Viscous Vortex Particle Method (VVPM) is utilized to generate rotor–rotor interference data for the Chopper Mars Helicopter platform, a large-scale hexacopter concept designed to be capable of carrying payload and pursuing independent science tasks. A reduced-order model compatible with finite state dynamic inflow is derived from the database. Interpolation strategies for continuous look-up are evaluated, with Gaussian Process Regression providing up to 20% improvement in prediction accuracy over linear interpolation of the interference data, although its scalability is limited by the large number of output channels. The interference model is implemented in HeliCAT, the flight dynamics analysis framework used for the Ingenuity Mars Helicopter, to assess the impact of rotor–rotor interference on trim, open-loop response and key stability derivatives. For Chopper, the first-order impact of rotor–rotor interaction on flight dynamics in hover is small, but with an observed increased pitch and roll damping. In forward flight, on-axis bare airframe responses exhibit attenuation at the lower frequencies due to interference, and increased sensitivity of the pitch response to vertical perturbations in cruise. Finally, a real-time closed-loop simulation is performed with and without the interference model to assess the impact on rotor power during a representative science mission flight trajectory on Mars, showing that total increase in mechanical shaft power due to multirotor wake interactions does not exceed 5% throughout the flight.
A high-fidelity computational study is conducted to investigate the aerodynamic behavior and flight response of an electric Vertical Take-Off and Landing (eVTOL) multirotor configuration using unsteady computational fluid dynamics (CFD) framework. Four simulation cases are considered to examine the vehicle aerodynamics under both prescribed and fully coupled conditions. Prescribed hover and forward-flight cases isolate rotor aerodynamics and rotor-airframe interactions under constrained kinematics. Six-degree-of-freedom (6-DoF) free-flight maneuvering simulations capture the coupled evolution of aerodynamic loads, vehicle attitude, and translational motion. The results demonstrate that the high-fidelity unsteady CFD framework, coupled with rigid-body dynamics, effectively resolves the tightly coupled aerodynamic–dynamic interactions inherent to eVTOL configurations. This work provides a foundation for future investigations into trim strategies, control modeling, and expanded flight envelopes.
A 4.75-ft diameter hingeless hub proprotor model was wind tunnel tested up to the very high speeds of 205 knots, loosely corresponding to 480 knots full-scale, with parametric variations in blades, wing spar, and pylon center of gravity. Testing revealed that a gimballed-hub configuration that reached whirl flutter at 160 knots was completely stabilized when converted to a hingeless hub – using identical blades, span, and pylon. While the gimballed-hub model encountered whirl flutter at 160 knots, the hingeless-hub configuration remained stable throughout the entire test envelope up to 205 knots. The key conclusions are that a hingeless hub can eliminate whirl flutter, and that the most stable configuration is a swept-tip blade hingeless-hub rotor with the pylon center of gravity aft of the wing spar.
This paper discusses the development of a quantitatively-accurate non-linear hybrid flight dynamics model of a hover-capable Air-Launched Tailsitter Unmanned Aerial System (ALUAS) in order to 1) understand its dynamics during complicated maneuvers, and 2) provide a high-fidelity framework to develop novel control laws. Wind tunnel tests were conducted on a 1:1 scale model of the full aircraft to measure the airloads, which were used in the simulation as a lookup table. Flight tests of the ALUAS were performed in hover, transition, and cruise to collect a large amount of unique state measurements by providing large excitations to induce highly transient motion. The flight dynamics predictions using Rotorcraft Comprehensive Analysis System (RCAS) software were then compared with experimental flight test data. To correct any discrepancies in the RCAS physics-based predictions, a correction was learned from the experimental measurements, making use of the large amount of collected flight test data. Using a neural network to learn this correction, the end result was a quantitatively accurate neural network assisted flight dynamics model. The accuracy of current simulations in complex flight states successfully demonstrates the applicability of the proposed methodology for correcting the dynamics model of novel out-of-the-box aircraft configurations.
This study introduces a structured methodology for identifying Control-Equivalent Turbulence Input (CETI) models using rotorcraft flight dynamics simulations. A new Moving Spatial Turbulence Field (MSTF) model was developed to generate input datasets, enabling CETI model identification for four distinct aircraft configurations: a generic utility helicopter resembling the H-60, and three small-scale multi-rotor UAS types—a quadcopter, hexacopter, and octocopter. The CETI models were validated in hover using frequency-domain analysis, with flight-derived CETI models serving as the benchmark. To further assess model performance in forward flight, CETI models for the H-60 were identified at airspeeds ranging from 0 to 140 knots in 40- knot increments. Results indicated that the MSTF-based CETI models for the H-60 effectively captured key spectral features of the flight-test data, though some deviations were observed, potentially due to variability in atmospheric conditions. In contrast, the CETI models for the multi-rotor UAS configurations showed closer alignment with flight data, likely reflecting more accurately captured atmospheric inputs during simulation. These findings highlight the promise of simulation-based CETI modeling for both rotorcraft and UAS, offering valuable insights into turbulence effects across diverse platforms and flight regimes.
This paper details the development of a tailsitter unmanned aerial system (UAS) that has the potential to be airlaunched in the near future. By simultaneously integrating air-launch capability with both rotary-wing vertical flight and fixed-wing horizontal flight, the vehicle can be rapidly deployed, perform hovering flight, and achieve high-speed and efficient cruising flight. The aircraft prototype has a mass of 1 kg (2.2 lbs) with wings that can fold to allow the aircraft to fit inside a 6-inch launch tube. A coaxial propeller with vectored thrust is used for control in vertical flight, and a unique avian-inspired wing-folding mechanism is used for stowing and deploying the wings. The aerodynamic design was characterized through a series of wind tunnel experiments, propeller tests, and flight dynamics simulations. High-fidelity simulations of vehicle dynamics validated its air-launch capability and flight tests performed with the prototype demonstrated the ability of the aircraft to perform hovering, transitioning, and fixed-wing cruising flight. To date, this aircraft has achieved speeds of up to 130 km/hr (81 mph), and range up to 10 km (6.2 miles).
A piloted simulation study in the Vertical Motion Simulator at NASA Ames Research Center will investigate the handling and ride qualities of eVTOL configurations (lift-plus-cruise and tiltwing) for both civilian and military applications. The flight dynamics models were developed in the FLIGHTLAB modeling and analysis software environment, while explicit model-following control laws and high-fidelity powertrain models were developed in Simulink. The Joint Input-Output method was used to generate frequency responses for linear model verification, as the control effectors are highly correlated for these types of vehicles. The linear models were verified for the frequency range of interest for handling qualities. Once verified and tested individually, the three parts (flight dynamics model, control laws, and powertrain) will be integrated into the Vertical Motion Simulator for piloted simulation evaluations.
This paper describes the dynamic modeling and flight control software development efforts for a subscale tiltrotor electric vertical takeoff and landing (eVTOL) aircraft built at NASA Langley Research Center. The vehicle, referred to as the Research Aircraft for eVTOL Enabling techNologies (RAVEN) SubscaleWind-Tunnel and Flight Test (SWFT) model, serves as a flight dynamics and controls research testbed to foster advances in eVTOL aircraft technology. After fabricating the vehicle, wind-tunnel testing was conducted to identify a high-fidelity aero-propulsive model for use in a flight dynamics simulation enabling flight control system development. The RAVEN-SWFT aircraft subsequently underwent flight-test risk reduction steps and then free flight testing employing custom research flight control software. The flight control software, which can be efficiently updated and tested on the vehicle, includes a robust model-based control algorithm and an extensive programmed test input injection capability. The progress of RAVEN-SWFT research activities will be summarized alongside the associated modeling and flight control software development aspects, including the flight dynamics simulation, flight control system architecture, vehicle integration, and testing approaches.
The next generation of Mars rotorcraft may involve an increase in scale and number of rotors. A key focus area that has been identified is to increase the fidelity of rotor wake modeling, including its impact on flight dynamics. To that end, this paper pursues the use of a Viscous Vortex Particle Method (VVPM) for mid-fidelity rotor wake predictions in Mars atmospheric conditions. Simulated aerodynamic hover performance, as well as control efforts in trimmed forward flight, of the Ingenuity Mars Helicopter with a VVPM wake is shown to correlate well with available experimental data. Qualitative and quantitative coaxial wake effects for Ingenuity-type rotors in hover and forward flight as predicted with VVPM are studied. Utilizing VVPM to evaluate rotor-rotor interference effects in a large-scale Mars hexacopter across a wide range of flight conditions showcases the capability to comprehensively model the induced wake of complex multi-rotor configurations within feasible computational cost. An approach to emulate interference effects derived from the VVPM data that is compatible with finite-state dynamic inflow models is demonstrated, exemplifying how VVPM can be adopted to inform flight dynamics analysis and control system design of next-generation Mars rotorcraft.
This paper discusses the development of a flight dynamics model (or digital twin) of a compact and re-configurable coaxial-propeller-based micro air vehicle (MAV) in hover, edgewise, and maneuvering flight using a hybrid physics-based plus data-driven approach. The MAV has a mass of 366 grams (0.81 lb), and features a 52 mm (2.05 in) diameter cylindrical fuselage, foldable propellers, and a two-axis gimbal thrust vectoring mechanism for pitch and roll control. The aircraft has been successfully launched from a pneumatic cannon and has achieved stable and controlled flight. A physics-based flight dynamics model of this novel MAV has been developed using Rotorcraft Comprehensive Analysis System (RCAS). RCAS is able to predict the translational dynamics near hover reasonably well; however, the accuracy decreases for rotational dynamics in edgewise flight resulting in significant differences between predicted dynamics and flight test data, known as residual dynamics. The current hybrid model utilizes the residual dynamics via a data-driven approach to correct the physics-based model. Using the measured vehicle states and control inputs, a deep neural network (DNNs) was trained to learn the residual forces and torques. The resulting hybrid model reduced prediction errors by 55% on average compared to the RCAS model based on pure physics.
Dragonfly is a rotary-wing lander, and its mission is to explore Titan. It will make multiple flights over several years to explore different sites on Titan. There is limited information on the chemical processes that led to life on earth. Among the other places in the solar system, Titan is the most like the early earth and therefore exploring its organic surface chemistry will help to better understand our own prebiotic history. During Titan flight the rotor induced unsteady aerodynamic loads, as well as the interactional aerodynamic loads due to the rotor to rotor and rotor to lander interferences drive the structural vibrations. Therefore, robust and accurate predictions of Dragonfly structural loads and vibrations are essential for designing a vehicle that can successfully perform its mission. This paper presents the structural loads and vibration predictions of the Dragonfly lander using Rotorcraft Comprehensive Analysis System (RCAS) coupled with the Viscous Vortex Particle Method (VVPM) inflow model as well as the high fidelity coupled Computational Fluid Dynamics/Computational Structural Dynamics (CFD/CSD) simulations. Elastic fuselage modal model, generated by the NASTRAN FEA, is also used to represent the dynamics of the flight lander in the simulations. Hub loads and the lander vibrations simulations are provided at multiple flight conditions. The paper also demonstrates how the comprehensive analysis can be used to provide insight on the design changes that can result in significant reduction of the structural loads and vibrations.
The advent of electric propulsion technology has led to a paradigm shift in aircraft design over the past few decades. This shift has expanded the possibilities for design and optimization processes more than at any previous time. To support these advancements, efficient flight dynamics simulation models that can be employed in iterative optimization and design processes are essential. Among the modules of a typical flight dynamics framework—namely, control, flight dynamics, and aerodynamics—the aerodynamics module, which includes the rotor performance model, generally demands the most computational effort, thereby limiting simulation efficiency. In this study, a novel machine learning (ML)-assisted flight dynamics framework is developed, incorporating a Neural Network Blade Element Theory (NN-BET) model as the rotor performance module. The results show a 7- to 8-fold reduction in computational time compared to fast, physics-based frameworks utilizing efficient Blade Element Momentum Theory (BEMT) models, without compromising predictive accuracy. Furthermore, the modular architecture of the framework allows for easy adaptation to a wide range of practical applications by replacing modules with functionally equivalent alternatives. The demonstrated accuracy and computational efficiency of the proposed flight dynamics framework make it a highly promising candidate for optimization and design applications.
This paper explores the effect of addition of a horizontal tail on the longitudinal stability and performance of a Biplane Tailsitter Unmanned Aerial Vehicle (UAV). Biplane tailsitters a type of hybrid UAVs, often exhibits poor longitudinal stability during forward flight, necessitating continuous active control through application of differential motor thrust to maintain attitude. To address this challenge, this work proposes the integration of a horizontal tail on a quadrotor biplane tailsitter UAV, aiming to improve pitch stability and control authority during critical flight phases. Experimental flight data was utilized to determine the appropriate sizing of the elevator. A detailed flight dynamics model validated the effectiveness of the elevator control. The design was validated through outdoor flight testing, comparing the performance of tail-less and tail-attached configurations. The results demonstrate that the modified design results in a reduction control power requirement for pitch control on the motors, enhancing longitudinal stability, and improved ascent and descent rates. This study provides a systematic approach to improving the operational capabilities of tailsitter UAVs, contributing valuable insights for the design of future UAVs requiring enhanced maneuverability and efficiency in VTOL and forward flight modes.
Flight test students must explore a wide range of helicopter dynamic responses to learn how to assess conditions ranging from good conditions operation to those approaching, or even experiencing, loss of control. To introduce this evaluation process, the Flight Test and Research Institute (IPEV) implemented a helicopter flight dynamics model. This model is stitched in the x-body velocity (u) and y-body velocity (v) to achieve more accurate simulation, combined with a Variable Stability Augmentation System to assess different conditions prior to experiencing them in real flight. The use of robust control, where a fixed controller is applied to flight control systems under various operating conditions, presents an alternative to the traditional gain scheduling technique commonly used in aeronautical systems. This paper explores the potential to reduce controller design complexity while evaluating the impact on the helicopter’s full flight envelope through quantitative analysis and handling qualities evaluation by piloted simulation according to ADS-33E-PRF standard.
This paper demonstrates methods of aircraft sizing, flight dynamics modeling, and performance analysis using a lift+cruise concept vehicle with an electric powertrain and variable-speed rotors. The central focus is the development of methods to relate the aircraft design sizing constraints to achievable maneuverability and predicted handling qualities. A toolchain is demonstrated that performs aircraft sizing, mass moment of inertia estimation, powertrain modeling, trim optimization, dynamics linearization, handling qualities prediction, and quantification of achievable maneuverability under both nominal conditions and control effector failures. A convex optimization problem framework is introduced to compute agility bound estimates without requiring control system design or control allocation, potentially supporting rapid design iteration as well as early detection of deficiencies and undesirable operating conditions. This analysis is supplemented with more conventional methods of analysis to provide additional perspective and observations. Overall, the results suggest that each modeling and analysis element in the demonstrated toolchain adds significant value, with the combined approach supporting a more efficient and comprehensive exploration of trade-offs within the design space.
Rotors and propellers in edgewise flight typically encounter reverse-flow on the retreating blade, especially when operating at low rotational speeds and high speed flight. This phenomenon is well known and has been observed in rotorcraft and vertical take-off and landing (VTOL) applications, with impacts on vehicle performance and aerodynamic loads. Reverse flow is characterized by flow incident to the trailing edge of an airfoil with an angle of attack (AoA) of around 180°. Aerodynamic coefficients for reverse flow conditions are difficult to find in literature, and wind tunnel measurements often focus on the normal operating range of airfoils. This study investigates the fundamental aerodynamic characteristics of airfoils in reverse flow using high fidelity computational fluid dynamics, and analyzes the impact of using accurate aerodynamic coefficients on comprehensive rotorcraft analysis. Although the effect on flight performance is well understood, for applications on lift rotors of eVTOL configurations, reverse flow phenomena are expected to have a significant impact on rotor loads, especially during flight transition from VTOL to fixed-wing modes.
This paper investigates the use of multi-modal cueing through full-body haptic feedback to enhance pilot-vehicle system (PVS) performance, reduce mental workload (MWL), and increase situational awareness (SA) in both good and degraded visual environments (GVE/DVE). Piloted simulations were conducted using an H-60-like flight dynamics model in a virtual reality (VR) motion-based simulator, evaluating two ADS-33-like mission task elements (MTEs) – precision hover and slalom – under visual-only and combined visual and haptic feedback conditions in both GVE and DVE. The H-60 flight dynamics were augmented with a dynamic inversion (DI)- based stability augmentation system (SAS), implementing rate-command/attitude hold (RCAH) response type on the roll, pitch, and yaw axes and altitude hold response type on the vertical axis. The SAS was designed to achieve Level 1 handling qualities per ADS-33 standards. The full-body haptic cueing strategy leveraged an outer-loop DI control law, which provided vibrotactile feedback to cue desired roll, pitch, and yaw attitudes to the pilot. Roll cues were delivered via tactors mounted on the upper arms, pitch cues via tactors on the chest and back, and yaw cues via tactors on the calves. Eight test subjects participated in the piloted simulations, including three U.S. Navy test pilots and five subjects with different flying experiences. Results indicated that haptic feedback significantly improved hover performance, reducing MWL and enhancing SA, particularly in DVE. However, in the slalom task, predefined haptic guidance misaligned with pilots’ individual control strategies, leading to performance degradation. This finding highlights the need for pilot-specific adaptive haptic feedback to mitigate inconsistencies in dynamic maneuvering tasks.
The use of sub-scale vehicles as a means of predicting full-scale vehicle behavior has historically been applied to flight dynamics testing and evaluation for aircraft operating in Earth atmospheric conditions. However, the use of sub-scale testing on Earth has not been as thoroughly explored for Martian rotorcraft. In this paper, sub-scale vehicles of varying sizes were developed in simulation using Froude scaling laws to evaluate their ability to estimate fullscale linear dynamics for the Mars hexacopter, Chopper. Blade loading, Lock number, and flap frequencies were held fixed when scaling and corresponding relationships for vehicle length, mass, inertia, and rotor speed derived. Full-scale frequency response, gain margin, and instability characteristics are explored for hover and forward flight cases in a variety of Mars-to-Mars and Earth-to-Mars conditions. Mach effects are also analyzed as a consequence of Froude-scaling by comparing sub-scale vehicles that are Mach-matched to the full-scale Chopper. Finally, a first-order approach for downselecting a sub-scale vehicle based on feasibility is introduced.
Precision flight in windy conditions is a common challenge for multirotor UAS. It is especially challenging for in contact tasks that require high-precision positioning and good disturbance rejection capabilities. Such tasks include landing on high-voltage powerlines for in-contact inspections. This paper presents the implementation of small lateral thrusters to improve the lateral position hold ability of a large power line inspection UAS in windy conditions. Arranged in antagonistic pairs on each side, the lateral thrusters handle the high-frequency but smaller-amplitude wind turbulence components with a frequency split control. Using an identified model of the UAS flight dynamics alongside flight data in high-wind conditions, a control architecture with a frequency split in the lateral axis was optimized to increase the disturbance rejection. Experimental tests showed a 67% reduction in lateral position error with the proposed approach in high-wind conditions.
T-625 Gökbey is a light utility helicopter developed by Turkish Aerospace Industries since 2013. For T-625, automatic flight control system performance evaluation and development test campaign was conducted. In this paper, test campaign is investigated thoroughly. To assess and quantify the automatic flight control system performance and handling qualities, various different metrics and specifications were selected. This metrics covered both time and frequency domains. After metric selection, a set of test points were created. Most of the test points required delicate piloting and were easy to fail. Furthermore, a large flight envelope in terms of altitude and air pseed was needed to be covered. Hence, both test point number and required flight time estimates were very large. Hence, to not further increase flight time and need of test point repetitions, various different precautions were taken, such as using computer generated sweeps. While conducting tests, altitude kept constant while speed envelope tested. Then, altitude were changed to a new value and speed envelope is tested again. After each sortie, flight data were analyzed whereas control parameter updates were performed after finishing tests for all airspeed values for an altitude. After control parameter updates, validation tests were done to ensure that updated parameters provided expected performance. At this point, most of the test campaign is done and selected specifications were met for almost all the tested flight envelope.
A system identification study was conducted on a quadrotor unmanned aerial system (UAS) that was free-flying inside the test section of the Naval Surface Warfare Center Carderock Division's Subsonic Wind Tunnel. Motion capture cameras installed in the wind tunnel provided position feedback information to the aircraft in real time, enabling autonomous flights. Longitudinal, lateral, heave, and yaw axis frequency sweeps were conducted at airspeeds up to 20 knots, in 5 knot increments. The extracted flight dynamics model showed excellent agreement in both the time and frequency domains across all airspeeds. Variation in the aircraft's stability derivatives, power usage, and trim information with airspeed was determined. This paper documents the test procedures, challenges with flying aircraft inside the wind tunnel, the controller model, and the system identification results. This free-flight wind tunnel testing methodology has wide applicability to assist with UAS flight control development, and represents a new opportunity in the way traditional wind tunnels can be utilized for UAS research, development, and testing.
Rotorcraft responses to idealized disturbances are examined to gain insights into model fidelity requirements for flight simulations of the ship-rotorcraft dynamic interface. Two disturbance fields are considered: an isolated straight vortex that represents the canonical vortex that results from the corners of flat top ships in oblique wind-over-deck conditions and a horseshoe vortex derived from a nondimensional characterization of the time-averaged flow observed aft of a simplified ship superstructure. Rotorcraft models considered include: an analytical blade element theory-based rotor model, where the disturbance velocities are integrated over the rotor, and a coupled blade element / free wake flight dynamic model of the full UH-60 aircraft, which is used to perform time-marching simulations with the disturbances modeled as a frozen field that is fixed in space and not interacting with the aircraft (one-way coupling), and as a distorting field (two-way coupling). Analytical thrust responses of four idealized rotors to the disturbance associated with the horseshoe vortex suggest the existence of a scaling relationship between nondimensional blade loading and the product of nondimensional vortex width and strength. Time-averaged rotor thrust responses to the disturbance induced by the isolated straight vortex indicate significant distortion of the rotor wake and disturbance field at low forward speeds, with the frozen field assumption indicating greater disk tilt responses. Modeling the horseshoe vortex disturbance as a frozen field generates a greater rotor response than in the distorting case. Inclusion of full aircraft dynamics attenuates rotor responses to the disturbances when compared to isolated rotor responses.
The advent of electric propulsion is revolutionizing the paradigm of rotorcraft design, leading to new electric Vertical Take-Off and Landing (eVTOL) aircraft. Direct drive topologies are common within these new designs, and some designers have chosen to utilize this mechanism for Primary Flight Control (PFC), effectively utilizing the aircraft engines as PFC actuators to control the speed of the rotors. This decision integrates the propulsion and flight control systems, and intrinsically couples the aircraft sizing and control. Four separate tools were exercised throughout this study to conduct a conceptual design exploration of eVTOL aircraft handling qualities. The main tasks for these tools were: 1) aircraft sizing and performance analysis, including the calculation of trim; 2) flight dynamics modeling and analysis; 3) handling qualities-centric control law optimization; and 4) electric motor sizing. Sizing of an RPM-controlled Hexacopter concept explored the dependency of aircraft size to fundamental design parameters: 1) disk loading and 2) blade loading coefficient. Increasing the design disk loading resulted in designs with high agility, but at the cost of significant growth in the design gross weight. Finally, Categories II and III pilot-induced-oscillation (PIO) are known potentially-catastrophic handling qualities deficiencies of fly-by-wire flight control systems such as those expected of eVTOL aircraft. Consideration of PIO predictive metrics in conceptual design control synthesis led to increased PIO robustness.
In this paper a full vehicle UH-60A helicopter with and without stub wings is simulated in steady autorotation using both the flight dynamics (FD) code HeliUM-A and the computational fluid dynamics/computational structural dynamics (CFD/CSD) solver CREATE-AVTM Helios. Significant effort was put into trimming the CFD/CSD simulations to an autorotative state that is representative to how the aircraft flies in practice, and the method for doing so is described. For the baseline aircraft, both FD and CFD/CSD predictions of the vehicle's autorotational descent rate were in excellent agreement with available flight test data. CFD/CSD analysis showed that the presence of the fuselage led to a significant positive pitching moment on the rotor as well as a sudden loss of loading as the rotor blades pass through the fuselage wake. When External Stores Support System (ESSS) stub wings were added to the model, both solvers show both quantitative and qualitative agreement in their predictions of the increase in autorotational descent rate due to the additional ESSS drag. CFD/CSD analysis showed that the stub wings had minimal effect on the interactional aerodynamics between the airframe and rotor.
This paper discusses the development, flight-testing, and flight dynamics modeling of a Micro Air Vehicle (MAV) that could be deployed in a folded configuration via hand launching. This 112-g MAV features folding propeller arms that can lock into a rectangular profile the size of a smartphone. The vehicle is designed to be rapidly deployable via simply throwing it in the air, at which point the arms would unfold and the vehicle would autonomously stabilize to a hovering state. The capability of the feedback controller to stabilize the MAV from different initial conditions including tumbling rates of up to 2500 deg/s was demonstrated via flight-testing. A six-degree-of-freedom flight dynamics model was developed and validated using flight test data obtained using a motion capture system for various hand-launched scenarios. The proposed MAV, with its extreme portability and rapid/robust deployment capability, could be ideal for emergency scenarios, where a standard launch procedure is not feasible.
Rotorcraft experience significant vibrations due to periodic aerodynamic forces and moments on the rotor blades and wings. Rotor torque damping is a novel vibration damping method which uses small torque perturbations from the main electric motor to reduce vibrations. The large inertial and aerodynamic rotor loading and relatively high frequency torque perturbations mean that the rotor speed changes are small, so the rotor thrust and flight control performance are not significantly affected. This paper investigates the application of electric motor torque control for damping structural vibrations of an aircraft. The structural dynamics of the aircraft are represented using a finite element model of a quad tiltrotor eVTOL. Using collocated angular rate feedback on all four rotors provides more than 10% damping in controllable modes. The RMS value of flap-wise angular rate can be reduced by 91% with less than 1.2 RPM rotor speed change in response to a 20% vertical step gust in airplane mode. For N/rev disturbance cancellation, an optimal controller is designed assuming known disturbance location and frequency for active vibration control (AVC). The transfer matrix of a single cantilevered wing is calculated and used to feed forward harmonic rotor torques. The wing undergoes aerodynamic disruptions at N/rev and the harmonic controller reduces N/rev shear force and bending moment at the root by 20% and 58% with less than 1 RPM rotor speed change, respectively.
Multirotor UAS spanning Groups 3 and 4 have received increased attention as candidates for tactical resupply missions due to their VTOL capability and payload capacity. The objective of this work is to better understand how the parameters of multicopter UAS flight dynamics models scale with size in support of expanding the Army's unmanned aerial reconnaissance capability. A family of coaxial multirotor UAS spanning Groups 2 and 3 have been flight tested to gather data for flight dynamics modeling and validation. These UAS consist of the TRV-80, TRV-150, and the subscale Eagle platform. A series of test points including static stability, trim shot, frequency sweeps, doublets, and maximum climb rate maneuvers were collected. Wind data was simultaneously collected using a 3-axis ultrasonic anemometer to characterize wind conditions and characteristics during testing. Flight data were collected in varying payload configurations ranging from 0-120 pounds and at flight conditions ranging from hover to 50 knots. Point state space models were identified from frequency sweeps and verified using doublet time histories. Point models of the TRV-150 were combined with trim data and experimentally measured moments of inertia to develop a stitched simulation model. The stitched model was used to extrapolate for the flight tested off-nominal payload configurations for analysis. The unstable roll and pitch hover eigenvalues between the vehicles within the TRV family and smaller Group 1-2 UAS were analyzed and compared to Froude-scaled results to establish trends based on vehicle aspect ratio.
Over the past decade, due in large part to heavy investment in the field of Advanced Air Mobility (AAM), significant progress in rotorcraft-focused modeling tools has been made. Such progress has notably increased AAM rotorcraft modeling capabilities in the topics of conceptual design, preliminary design, and more recently flight dynamics. Yet, due to recent and persistent increases in extreme weather events, an emerging interest has been raised in utilizing such modeling capabilities for aiding in emergency relief efforts and other public good missions. This paper uses wildfire fighting as a representative public good mission and demonstrates the relevance of the NASA Revolutionary Vertical Lift Technology (RVLT) rotorcraft toolchain to such missions. An emphasis is placed on flight dynamics modeling and control because of the hazards and challenges associated with the atmospheric environment of wildfires. In this work, the NASA FlightCODE tool was used to analyze both a UH-60 and the NASA six-passenger quadrotor reference model hovering in an experimentally informed wildfire turbulent environment. Preliminary results of this study estimate actuator usage exceedances and disturbance rejection capabilities of the vehicles' translational rate command systems. Leveraging the RVLT toolchain, refinement and expansion of this work could lead to handling qualities envelope estimation and design optimization for wildfire turbulent environments. This would provide pilots with additional information to make real-time decisions in high-risk scenarios and begins preparations for simulating these dangerous environments for pilot training and experimentation.
Many traditional ship-rotorcraft interactional simulation approaches, including those used for pilot training, use a one-way coupling between aerodynamics and flight dynamics. In a one-way coupled method, the standalone ship airwake is superimposed on the rotor, modifying its inflow. However, because the rotor wake does not alter the ship airwake in such a simulation, one-way coupling may not capture all relevant phenomena, such as dynamic ground and wall effects; two-way fully-coupled simulations may be needed. In this study, one- and two-way coupled realtime and near-real-time simulation models of the ship-rotorcraft problem were developed using a GPU-accelerated Lattice-Boltzmann Method (LBM) flow field solver. Comparing flow fields and rotor hub loads, the two-way coupled simulations showed good agreement with new ship-rotor experimental data from Georgia Tech. Real-time full-scale rotorcraft ship approach maneuvers of a notional UH-60A landing on the NATO Generic Destroyer were conducted using a flight controller, and the results showed significant differences in flight control inputs between one- and two-way coupling methods. On the other hand, the presence of ship motion was found to have little effect on the control inputs. The study proved the feasibility of real-time fully-coupled ship-rotorcraft flow field simulations, achieving real-time performance on a single GPU node, and it paves the way for fully-coupled LBM pilot-in-the-loop flight simulation trials.
Coupled powerplant and rotorcraft flight dynamics simulations are commonly carried out in the non-linear time-domain framework (e.g. for pilot-in-the-loop handling qualities assessments), although these integrated models are generally not fully accurate from drivetrain dynamics perspective. Nevertheless, there is interest to verify that usual assumption of decoupled torsional stability (including rigid drivetrain analysis) and aircraft rigid body stability is valid, and up to what extent. The process described in the paper entails the automatic assembly of relevant subsystems (bare aircraft flight dynamics, Flight Control System including fly-by-wire actuation, sensors, and Control Laws software, drivetrain dynamics, powerplant dynamics) state space matrices through a Company developed Matlab toolbox. The proposed approach is control system design oriented, i.e. it does not require detailed flexible multibody modelling of the entire aircraft including dynamic systems and it is a natural extension of the process generally carried out by Control Laws and Flight Mechanics discipline for rigid-body stability assessment. The paper specifically addresses, as case study, the coupled torsional dynamics analysis of Next Generation Civil Tiltrotor Technology Demonstrator (NGCTR-TD) in Helicopter configuration. The NGCTR-TD project is developed under the EU Clean Sky 2 program for defining novel technology for future civil Tiltrotor platforms. The proposed coupled torsional stability modelling process is described, and the main outcomes valid for NGCTR-TD trimmed in hover and near-hover conditions presented. It is shown that the proposed methodology provides consistent outputs against the original NGCTR-TD linear stability results (from individual flight dynamics and drive system models). Analysis of Modal Participation Factors is used to quantify the coupling existing between drive system and flight dynamics state vectors, in nominal and failure conditions (Inter Connect Drive Shaft failure).
Advanced rotorcraft configurations currently being considered for Future Vertical Lift and Advanced Air Mobility applications typically feature redundant control effectors, which bring new opportunities for control design, including the ability to re-allocate control in response to failure or damage. This paper presents the design of damage tolerant control (DTC) for a generic utility-class tiltrotor using redistributed pseudo-inverse control allocation with axis prioritization. The damage tolerant control was integrated into full flight envelope control laws, and tested in a piloted simulation where pilots attempted to land the tiltrotor after damage in a cruise flight condition. Overall the results showed that damage tolerant control resulted in improved survivability ratings for the most severe damage cases. This work was done in support of the Adaptive Digital Automated Pilotage Technology (ADAPTTM) program which aims to develop a flight control software package to take advantage of redundant controls to improve safety, survivability, and performance for advanced vertical takeoff and landing capable aircraft.
In this paper, an offline path planning module, which is capable of generating dynamically feasible 3D trajectories for a class of Vertical Takeoff and Landing (VTOL) vehicles is presented. Input to the module is a flight plan defined by a set of way-points and its output is twofold: first, it produces an improved flight plan introducing additional waypoints and speed changes based on the heuristics and dynamical constraints of the vehicle. This new plan facilitates the pilot by providing information on specific locations and changes of the original flight path. Second, it generates a set of reference points, which can be used as the initial set of inputs for an online reactive trajectory optimization algorithm. The proposed development is capable of processing both climbs and descents as well as both fly-by and flyover waypoints, and speed changes in between those way-points. The module was also designed to capture the pilot's perspective of an abstract way-point mission. NRC has implemented a nominal Nonlinear Model Predictive Control (NMPC) based online planner using the reference trajectories generated by the proposed offline planning module successfully. The generated improved mission was implemented on a simulator having Bell 412 flight dynamics and promising results were obtained.
Advanced Rotorcraft Technology (ART) and the NASA Ames Aeromechanics branch have jointly developed FLIGHTLAB® simulation models for Advanced Air Mobility (AAM) VTOL concept vehicles. The overarching purpose of the simulation model development is to establish a set of well defined reference vehicles for FLIGHTLAB users and the rotorcraft community. The ongoing research effort and enhancement of these AAM simulation models to fulfill the role of quality reference vehicles is this paper's focus. The content of this paper expands on the established characteristics of these AAM models in three primary areas. First, enhancement of the lift+cruise and tiltwing models with elastic airframe properties is discussed. The process of setting up the elastic airframe model in FLIGHTLAB, as well as the impacts on flight characteristics are explained. The introduction of the elastic airframe modeling allows these models to be used in flight dynamics, loads, and vibration analysis of the configuration designs. Next, linear model generation from the enhanced simulation model is covered. Confirming the validity of the linearized models is of importance, as these linear models are utilized for flight control design and tuning for these experimental configurations. For the final focus, the progress towards implementation of these models into the NASA Ames Vertical Motion Simulator (VMS) is described. This task seeks to demonstrate the procedures of integrating a FLIGHTLAB flight simulation in the VMS environment, test fully integrated simulation with communication between flight dynamics, control, and propulsion models, and explore the essential aspects of simulation model integration in a full flight simulator environment. This includes I/O definition, initialization, trim, flying, etc. By expanding the capabilities of the AAM simulation models, they continue to develop as valuable and approachable modeling references.
The Research Aircraft for eVTOL Enabling TechNologies (RAVEN) Subscale Wind-Tunnel and Flight Test (SWFT) model is a subscale aircraft built for flight dynamics and controls research demonstrated in wind-tunnel and flight-test experiments. The intent of this paper is to provide a summary of past, current, and future efforts being pursued by the RAVEN-SWFT project. Initially, vehicle development guidelines were crafted by a multidisciplinary team to ensure that the RAVEN-SWFT vehicle was well suited for research in multiple areas, including aero-propulsive modeling, flight controls, and autonomy, among others. The vehicle has been used to obtain extensive wind-tunnel data, enabling aero-propulsive model development across the transition flight envelope and validation of computational tools. The vehicle will be used to conduct flight testing in order to evaluate modeling strategies and flight control logic. The RAVEN-SWFT model also serves as a risk reduction activity for a conceptual, full-scale vehicle in the 1000-lb class. The next steps in the project are to successfully demonstrate free flight in hover, transition, forward flight, and the reverse thereof, utilizing custom control laws integrated onto the RAVEN-SWFT avionics hardware. The project intends to publicize all of the geometry, data, and methods in future reports.
A quadrotor was modified by adding wings to the frame to directly compare the flight dynamics characteristics as well as the stability and control derivatives of the quadrotor and its biplane tailsitter variant. The on axis response of the quadrotor and a biplane tailsitter variant were measured through flight test and frequency domain system identification was used for non-parametric and parametric model identification. Identification of the full vehicle dynamics demonstrated that also identifying the motor torque and back-EMF constants from no-load measurements and the remaining motor parameters from a rotor-motor test stand provided the most accurate identified full vehicle model. The motor dynamics were shown to add a pole to the thrust-based responses (roll, pitch, and heave), while the torque based response (yaw) included a pole and a zero. This approach was then used to identify and compare the quadrotor dynamics, tailsitter dynamics, and the total impact of canting the motors. It was found that the presence of the wing added pitch damping to the dynamics and pitch stability became negative. The yaw axis saw an increase in yaw damping derivative, and a reduction in the yaw control derivative to the point where it became difficult to control the aircraft. By introducing cant, both the quadrotor and tailsitter saw large increases in the yaw control derivative. Further, the rotor thrust based moment generation due to cant resulted in the yaw response zero being canceled by the motor dynamics, resulting in a purely first order yaw response. Neither the wing nor cant produced any change in the lateral and heave axes.
This paper addresses the urgent need to enhance rotorcraft safety and performance by developing a prediction methodology for the onset of the Vortex Ring State (VRS), and therefore verifying the VRS avoidance diagram. The objectives of this research are to assess the correlation between predictions generated by a comprehensive flight dynamics code and the latest and most accurate VRS boundary models, validate the VRS avoidance diagram across diverse descending flight conditions, and identify specific parameters indicating the rotor's entry into the VRS. The methodology involves a detailed investigation of 8 descent manoeuvres using a comprehensive flight dynamics code coupled with an advanced free vortex wake model. Results show that the pitch and roll oscillations and thrust fluctuations experienced by helicopters during the VRS are also observed in the model response to steep descent maneuvers. The findings confirm the reliability and applicability of the VRS avoidance diagram, providing valuable insights for improving rotorcraft design and flight control systems. Furthermore, this study presents novel findings regarding the identification of the VRS onset boundary through the observation of pitch and roll rate fluctuations, which contribute to the systematic detection and prevention of the VRS.
Pilot models have been successfully utilized for design and analysis of rotorcraft for decades. These models are commonly used as analysis tools, usually coupled to flight dynamics models. A method that has been developed is the Task-Pilot-Vehicle (TPV) approach, which utilizes the pilot model in a specific task. This paper presents results from the application of the TPV approach to investigate envelope protection (EP) methods. Four EP methods were chosen for these studies, i) Control Limiting, ii) Command Limiting, iii) MPC Based, iv) Virtual Control Limiting. These methods are exercised with pilot models inside a specifically designed toolbox across several different vehicles, including representative Tiltwing and Quadrotor vehicles, and Handling Qualities Task Elements (HQTEs). Results show that each envelope protection method has advantages and disadvantages, ranging from ease of implementation to potential adverse interactions with the pilot-vehicle system. Overall, the TPV approach shows promise as a method to investigate EP methods for eVTOL/VTOL aircraft.
This paper describes the design and initial flight testing of a compound coaxial tilting head rotorcraft (CCT-HR). Control is provided by titling the rotor head for roll and pitch, differential rotor speed for yaw rate, and rotor speed for total thrust. In addition, a longitudinal thruster is incorporated to enable higher speed forward flight and to add a degree of freedom for longitudinal trim in forward flight. The intent is to explore the feasibility of this vehicle concept and to develop a vehicle that can be used to explore control strategies. The steady state flight envelope is developed analytically; a simulation of longitudinal degrees of freedom is described and a control method for forward flight that incorporates the thruster is proposed. Results of near-hover flight tests are described and initial tests of forward flight using the thruster are described. The vehicle is shown to be stable and easily controllable near hover; in thruster-powered forward flight unmodeled rotor dynamics result in instability.
High-speed configurations are among the new emerging concepts that are currently expanding the scope of rotorcraft design. Especially in the field of defense technology research, the capability of a substantial increase in maximum velocity becomes more interesting. For instance, NATO project NGRC is considering this capability for a new medium utility rotorcraft. DLR supports these activities by its continuing defense technology research. In this study the benefits and drawbacks of a high-speed capability for a given mission scenario are analyzed. For that purpose, a contemporary configuration has been modeled, featuring a maximum velocity of 82 m/s (160 KTAS). The high-speed configuration meets with the same mission requirements, but with an increase of about 50% of maximum speed to 125 m/s (242 KTAS). All tasks in this study are conducted with DLRs integrated design environment IRIS. The high-speed configuration features an off-loaded main rotor, a wing, a propeller and a reduction of the rotational speed of the drive train for the extended flight envelope. The reference configuration as well as the high-speed rotorcraft were compared with the focus on flight performance.
Rapid advances in high fidelity modeling and high performance computing capabilities have enabled their routine utilization in support of aircraft design. Analysts are able to generate orders of magnitude more data that must then be turned into actionable intelligence to guide design. Enabling effective application of advanced analysis to design requires a robust end-to-end digital transformation to make the simulation processes reusable, repeatable, traceable, scalable and minimize setup errors. This is achieved through the development of a Computational Fluid Dynamic (CFD) modeling framework where streamlining and automation are inserted within the current CFD workflow that involves model setup, simulation and post processing. Workflow automation techniques have been implemented in simulation pre and post processing that reduce the overall process time or enhance the fidelity of the simulation. To conduct CFD evaluations through flight envelope efficiently, space filling methods that take into account uncertainties of complex systems are needed and have driven updates to the boundary condition and design of experiments (DOE) generation within the workflow. Vehicle sub-system design can be highly iterative, performed by a large number of participants in multidisciplinary groups. To ensure traceability across the digital thread, a provenance and metadata storage methodology has been implemented to capture information about CFD simulations and construct a query able database while a model-based systems engineering (MBSE) framework provides a structured and integrated approach to managing information throughout the product lifecycle. The SIM-FIX-SIM approach enabled with a robust analysis framework for digital flight assessment prior to first flight will contribute to the overall goal of reducing development timelines and achieving cost reduction goals for cutting-edge rotorcraft development programs.
This paper presents a real-time closed-loop rotorcraft simulation framework using HeliUM-A, a high-fidelity flight dynamics analysis, and a Simulink®-based flight control system model. Serial optimization and parallel computing techniques are introduced in HeliUM-A to achieve real-time speeds. A customized ordinary differential equation solver with parallel load balancing enables accelerated time marching simulations. Software interfaces are introduced to encapsulate HeliUM-A into a Level-2 S-function Simulink® block. Using standardized Simulink® ports, control inputs, rotor/body states and their time derivatives as well as relevant output quantities are communicated in-memory between Simulink® and HeliUM-A for closed-loop execution. This encapsulation retains the parallel computing improvements in HeliUM-A when executed through MATLAB, Simulink® or through the compiled executable automatically generated by the Simulink Coder. The framework is demonstrated on a coaxial compound scout helicopter with a pusher-propeller. Closed-loop vehicle and rotor responses are compared between two flight dynamics models: a stitched simulation model based on linearized models extracted from HeliUM-A, and the original HeliUM-A model. Time domain and frequency sweep comparisons showed excellent agreement between the stitched and nonlinear HeliUM-A models demonstrating that the new HeliUM-A based closed-loop framework shows expected behaviour. The HeliUM-A bare airframe model can be used to simulate various rotorcraft configurations in edge-of-the-envelope flight conditions, whereas the stitched models are accurate for center-of-the-envelope conditions.
This paper discusses the development of a fully-nonlinear flight dynamics model of a hover-capable Air-Launched Uncrewed Aerial System (ALUAS) in order to (1) understand the dynamics, controllability, and air loads of these type of aircraft while performing complicated maneuvers, (2) formulate design principles to feed back into the development of the realized physical aircraft, and (3) provide a high-fidelity dynamic framework to develop novel control laws. The flight dynamics model is developed using a software called Rotorcraft Comprehensive Analysis System (RCAS), where each component of the vehicle was modeled with varying fidelity. Wind tunnel tests were conducted on fullscale models to measure the forces and moments on the propeller, the isolated fuselage, and the full aircraft. Wind tunnel tests were also conducted to measure the forces and moments on the full aircraft for different wing folding angles. The thrust and torque of the propeller as well as the lift predictions for the aircraft correlated well with the test data while the drag and pitching moment were under-predicted. The qualitative trend of the forces and moments on the aircraft during wing unfolding were also correctly predicted. The RCAS model, once corrected with the test data, was used to simulate nominal maneuvers including edgewise flight, hover-to-cruise transition, cruise-to-hover transition, ground-launch, air-launch from a moving helicopter, and air-launch with asymmetric wing unfolding. These simulations demonstrated the capability of the vehicle to perform complicated maneuvers to meet various mission objectives in challenging environments.
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