Browse Topic: Flight dynamics
Dynamic responses at critical locations of a spacecraft due to excitations expected during the ascent phase of a launch vehicle mission are usually estimated through a Coupled Loads Analysis (CLA) using the structural dynamic finite element model of the launch vehicle coupled with that of the spacecraft. Generally, the full physical structural dynamic model of a spacecraft has lakhs of degrees-of-freedom (DOFs). Coupling such a model with a similar model for the launch vehicle results in exorbitantly high computational costs for CLA. Hence, dynamic analysis of such large and complex structural assemblies usually employ sub-structure coupling or Component Mode Synthesis (CMS) methods. The most widely used CMS method for dynamic analyses is the Craig-Bampton (CB) method. Conventionally, a full launch vehicle CLA involves one level of CB-reduction wherein a reduced-order dynamic model of the spacecraft is first generated using the fixed-interface CB-method. This reduced-order model is coupled with the launch vehicle model through the interface DOFs and CLA is performed using this coupled dynamic model. For test missions of ISRO’s manned space program, a simulated Crew Module (CM) is interfaced to the launch vehicle in place of the spacecraft. For CLA of this launch vehicle, a CB-reduced model of the CM is required. The CM comprises several sub-systems, including the Crew Seat Assembly (CSA), and dynamic responses at critical locations of these sub-systems need to be obtained from CLA. Structural dynamic model of the CSA is complex and involves lakhs of DOFs. Interfacing this detailed finite element model of CSA directly with the full model of CM for CB-reduction was found infeasible considering the prohibitively high computational time and resources required. An alternate approach is utilized to overcome the problem wherein the reduced-order dynamic model of CM is generated with two levels of nested sub-structuring. The CSA is represented as a CB-reduced model within the full physical structural dynamic model of CM, which is further CB-reduced and coupled with the launch-vehicle model for CLA. This paper presents the approach adopted to extract internal point dynamic responses on CSA from a dynamic analysis using the reduced-order dynamic model of CM with two-level CB-reduction. To validate the proposed approach, structural dynamic model of a skeletal structure of CM is generated encompassing the full model of CSA. A reduced-order dynamic model of this skeletal structure is also generated with two-level CB-reduction. A typical transient force excitation is considered and dynamic responses at internal points of CSA are estimated from the reduced-order dynamic model using the proposed approach. These responses are compared with corresponding responses obtained from a similar analysis with the full physical model of the skeletal structure and validity of the approach is established. The presented approach is generic and can be conveniently extended to extract responses from any dynamic analysis with models having even more than two levels of nested sub-structuring. Using this approach, the computational time and resource requirements for dynamic analysis studies are minimized. Dynamic analysis capabilities of MSC Nastran software are used for this study.
This paper investigates amplitude effects in the aeroelastic damping and frequency characteristics of the Maryland Tiltrotor Rig across four configurations: gimballed or hingeless hubs, each paired with straight or swept-tip blades. The recovery rate method is used to identify the aeroelastic parameters of the primary modes dominated by out-of-plane and in-plane wing bending from experimental free-decay strain time histories, capturing variations in dynamic behavior with the response amplitude. Results from conventional methods that assume linear (amplitude-independent) behavior are also presented for comparison. The local damping ratio of the examined modes generally decreases with increasing strain amplitude across all configurations, a trend missed by conventional linear estimation methods. The strength of amplitude effects varies as the system approaches instability: for gimballed configurations, they weaken near instability; for hingeless configurations, they become more pronounced. While the local frequency of the mode dominated by out-of-plane wing bending remains relatively constant with strain amplitude, the frequency of the mode dominated by in-plane wing bending displays significant amplitude-dependent shifts, particularly for hingeless hubs. The findings demonstrate the importance of accounting for nonlinear effects in aeroelastic parameter identification based on experimental tiltrotor data and provide insights into tiltrotor nonlinear dynamics.
The TiltRotor Aeroelastic Stability Testbed (TRAST) was developed to experimentally investigate whirl-flutter stability of tiltrotor aircraft. Previous wind-tunnel testing focused on configurations representative of current generation tiltrotors utilizing gimballed rotor hubs. The TRAST platform was also designed to support a hingeless rotor system to investigate whirl-flutter mechanisms representative of stiff proprotor configurations. This paper presents analytical whirl-flutter predictions for a hingeless rotor configuration of the TRAST model. Structural mode shapes derived from a NASTRAN finite-element model are combined with comprehensive aeroelastic analyses in CAMRAD II and RCAS. The results show that the dominant whirl-flutter mechanism differs from the gimballed configuration, with instability occurring through the wing in-plane mode rather than the wing vertical bending mode. Parametric studies examining rotor speed, pitch-spring stiffness, rotor flexibility, and diaphragm spring stiffness are conducted to evaluate the sensitivity of the predicted stability boundary. Results indicate that the hingeless configuration is significantly more stable than the equivalent gimballed configuration and exhibits different trends with rotor speed and structural stiffness. These predictions help identify configurations of interest for future wind-tunnel testing and provide insight into whirl-flutter mechanisms for hingeless tiltrotor systems.
Developing a comprehensive autonomy solution for the Army's current and future aircraft fleet requires a robust computational and perception capability for decision-making across the entire flight envelope without a pilot. This also requires a flight control system and infrastructure capable of executing autonomous decisions in complex mission environments. Ongoing development of automation and autonomy, utilizing a wide range of perception sensors, has been conducted on platforms such as Sikorsky's S-70 and the Army's UH-60Mx aircraft. This work builds upon previous efforts and leverages ongoing collaborations with industry, the Department of War (DoW), and the Defense Advanced Research Projects Agency (DARPA) to advance autonomous capabilities for both optionally piloted and uncrewed aircraft.
This paper presents control and estimation approaches for multiple vehicles to cooperatively sample atmospheric variables, with a focus on wind estimation, in complex environments. The technology is encapsulated in the WINDSENSE system. The collected data could be used to initialize weather models for nowcasting or forecasting, assess the fidelity of a meteorological model, assist in understanding plume dynamics or tracking plumes, provide real-time data for fire controls or wildfire fighting, or locate the source of a chemical, biological, radiological, or nuclear (CBRN) source. The wind estimation approach utilizes a Bayesian formulation with a process model found using system identification techniques. The model structure of the identified dynamics builds on prior work by the authors and combines first-principles and experimental data collection to generate a model that is valid over a wide range of the flight envelope. This enables the wind estimator to also be viable over the modeled domain, allowing for wind estimation at trim and quasi-trim conditions and during dynamic maneuvers. Performance of the wind estimator is validated in field testing via comparison with a research-grade ultrasonic anemometer. A decentralized controller enables coordinated flight amongst multiple vehicles to maintain a desired formation geometry and reshape the formation to optimize data collection for different environments or mission objectives. Sampling strategies for a swarm of vehicles are developed through analysis of large-eddy simulation studies of wind and particle dispersion using the PALM facility.
A new approach that enables the synthesis of fully coupled system dynamics is described in this paper. The approach facilitates collaboration between solver developers by explicitly avoiding inter-code coupling and instead uses a generic interface that enables inputs to be set and flags outputs available to other solvers. The assembly of a fully coupled linearized system matrix is obtained entirely from the existence of coupling maps, and does not rely on user intervention. Tiltrotor whirl flutter in cruise conditions is systematically investigated by careful examination of results obtained through various combinations of domain synthesis and obtained from the Hermes coupling framework. Investigated solver domains include nonlinear rotor dynamics, nonlinear aerodynamics, and linear structural dynamics. Utilized software modules include RCAS; Project Chrono, a purpose-built lifting line aerodynamics solver; and FuselageSolver for linear structural dynamics. Aeroelastic predictions are synthesized and verified for a pitch/plunge airfoil and a low-speed wing. Rotor aeroelastic effects are verified by coupling rotor dynamics with rotor aerodynamics. Structure-to-structure results are synthesized by coupling both RCAS and Project Chrono rotor dynamics models to a pylon structure modeled in FuselageSolver. Finally, full whirl flutter predictions are synthesized by coupling Project Chrono, lifting line aerodynamics, and FuselageSolver, as well as an RCAS model without the wing/pylon structure to FuselageSolver. Results indicate that the new synthesis approach is viable and accurate.
This paper introduces an eigenvalue-based whirl flutter prediction method accounting for aerodynamic interactions between a wing and propeller. The linearized unsteady vortex lattice method was utilized to model fixed-wing aerodynamics while the linearized viscous vortex particle method was utilized to model rotary-wing aerodynamics. The complete aerodynamics model was then coupled with computational structural models to demonstrate the capabilities of the model to predict whirl flutter using an eigenvalue-based method. Two computational structural models were used: the first being an analytical propeller model affixed to a rigid wing via root springs and dampers, and the second being the University of Michigan's Nonlinear Aeroelastic Simulation Toolbox. These models demonstrate the capabilities of the linearized aerodynamics model in predicting instability with structural models of different fidelities, both considering and not considering aerodynamic interactions. The linearized aerodynamics model predicts a reasonable aeroelastic solution when coupled with the structural models, but requires more investigation as to whether aerodynamic interactions are being sufficiently captured.
A 4.75-ft (1.45-m) diameter, dynamically-scaled proprotor with swept-tip blades was tested up to very high speeds of 205-kt (380-km/h) including the onset of whirl-flutter. Three important parameters that are difficult to vary at full-scale: hingeless hub, pylon placement, and wing spar, were examined consistent with both straight and swept-tip blades. The stability of all three wing-pylon modes: beam, chord, and torsion were measured. The in-house comprehensive analysis UMARC-II was used judiciously to shed light on the fundamental mechanisms at play and for validation. The key conclusions were: (1) on a gimballed hub, the swept-tip blade has no adverse effect on whirl-flutter, nor does it impede the mechanisms that might eliminate it, such as an aft pylon center of gravity placement or stiffer wing spar, and (2) on a hingeless hub, the swept-tip blade left the beam mode unaffected, but increased the chord>and torsion damping significantly through their interaction with the low-frequency regressive lag mode. Overall, the results demonstrate that swept-tip blades are a beneficial choice for high-speed flight when paired with an aft pylon, hingeless hub, or a combination thereof.
This paper utilizes a combined experimental and modeling approach to investigate techniques for improving the forward-flight roll-control authority of a Quadrotor Biplane Tailsitter (QBiT). QBiT is a mechanically simple, efficient hover/cruise aircraft whose roll authority in forward flight is traditionally limited by differential propeller-torque-based control. The two roll-control enhancement techniques investigated are propeller canting and the use of ailerons. A 2-kg instrumented QBiT platform was developed and flight tested to collect high-fidelity flight data across multiple flight regimes including hover, transition, cruise, and coordinated turns. A flight dynamics model was developed and validated using wind tunnel measurements and flight-test data. Flight tests showed that the cant-only configuration exhibited limited roll authority during coordinated turns due to motor control saturation, whereas the cant-plus-aileron configuration provided improved roll performance. Using test data from forward-flight roll excitation maneuvers, roll-control authority was evaluated both in the time domain and frequency-domain. The results showed that adding ailerons increased forward-flight roll-control authority by about 2.3 times from analyzing the flight data and by up to 2.6 times based on the flight dynamics simulations.
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.
Full state feedback offers theoretically guaranteed multi-axis stability, making it superior to conventional PID controllers. There is however one drawback, a full state controller has a mathematical difficulty if the B matrix is not square and thus not invertible. This is the case for helicopters with 6 degrees of freedom and 4 inceptors. Variations of linear quadratic regulators are a work around, however complexity dramatically increases. Best would be a direct solution to the original problem. This is the breakthrough result of this paper. This paper documents an approach which removes the analysis roadblock by partitioning the 6 x 6 system "A" matrix into two groups of 4 x 4 matrices. The 4x4 matrices are individually stabilized with full state gain matrices. One matrix is designated “Driver Matrix” which provides actuator commands. The other matrix is designated "Reference Matrix" which provides references. The two matrices are coupled together by requiring that the driver matrix follow references generated by the reference matrix. With each matrix individually stabilized, the coupled combination is also stabilized. Computation of flight dynamics states (u, v, w, p, q, r) is shared between the matrices. Initial results are very encouraging, showing an originally sluggish, heavy lift helicopter having now concise decoupled responses to pitch and roll commands. Stability derivatives are recomputed during flight allowing coverage over the whole flight envelope. A handling qualities task has been defined to relocate a 40 ft standard seaborne container directed by a pilot in a ground control station. Cooper Harper ratings of this task have demonstrated favorable Level 1 handling qualities if use is made of an automated lateral repositioning command.
This study investigates the dynamics and associated vibratory loads of an underactuated swashplate-less rotor and its impact on the flight dynamics of a small-scale helicopter powered by this rotor via a combined experimental and computational approach. Unlike prescribing cyclic pitch using a swashplate, here the pitch is a response to the 1/rev cyclic rotor speed input. This is enabled on the current two-bladed rotor using a skewed lag hinge that utilizes the cyclic speed variation to produce lagging motion and subsequently pitching the blades in a cyclic fashion (ƍ4 coupling) for generating the pitch and roll control moments. One of the key dynamic characteristics that distinguishes this rotor from a conventional swashplate-controlled rotor is that the two blades have dissimilar pitch and flap responses leading to high fixed-frame vibratory loads. Results show that a large 1/rev vertical shear force is transferred to the fuselage resulting in half-peak-to-peak loads of +/−0.67g. The flap responses of the two blades being out of phase caused a net inertial force, which was the dominant contributor to the vibratory vertical shear force. Upon removal of the flap hinge, the vertical shear force was reduced by almost 85% while increasing the control moment authority by 50%. The inertial moment about the rotor axis generated by cyclic lagging nearly balances the moment due to the angular acceleration (Ω ) of the rotor, which significantly reduced the dynamic torque requirement from the motor.
This paper investigates the impact of aerodynamic interactions on the dynamic aeroelastic stability of a wing-propeller configuration, with emphasis on whirl flutter. The wing structural dynamics are modeled using linear Euler-Bernoulli beam finite elements, while the propeller is represented using Reed's two-degree-of-freedom model. Baseline stability analyses neglecting aerodynamic interactions employ strip theory for the wing and the Houbolt-Reed formulation for the propeller. Analyses that account for aerodynamic interactions are then performed by coupling the wing and propeller structural models with the unsteady vortex-lattice method. Whirl flutter points are identified from transient simulations under both thrusting and windmilling conditions. Results show that three-dimensional aerodynamic effects increase the whirl flutter speed, whereas wing-propeller aerodynamic interactions play a slightly destabilizing role. Thrusting conditions produce a lower critical speed than the wind-milling case. The results demonstrate the viability of the unsteady vortex-lattice method as a unified aerodynamic framework for aeroelastic stability analysis of wing-propeller systems with mutual aerodynamic interactions. In addition, they reinforce findings from previous work that highlighted the destabilizing role of wing-propeller aerodynamic interactions.
Prior work demonstrated that acceleration washout in motion simulators produces decay-rate sensing ambiguity within the vestibular system, forcing pilots to rely on visual cues for control. While Pilot Induced Oscillation Ratings (PIORs) for flight and simulation have been matched using different sensing thresholds, a quantitative basis for the 50% reduction in the visual decay-rate threshold has remained elusive. This paper provides evidence that pilots perceive decay rate proprioceptively through stick force during both flight and simulation, rather than through vestibular or visual channels. The residues of the stick-force sensitivity transfer function reflect the amplification or attenuation of neighboring zeros and poles; when these residues fall outside the human's 30 dB tactile sensory window, the resulting decay rate becomes imperceptible. Modeling reveals that stabilization via the visual channel in simulators produces dominant mode characteristics - decay rates, frequencies, and residues - that diverge significantly from vestibular-stabilized flight. The Virtual Vestibular Cueing (VVC) technique is introduced to tune the visual channel using pseudo-vestibular rate cueing, optimally aligning the simulated dominant mode with flight-validated residues and decay rates. A preliminary fixed-base study demonstrates that VVC synchronizes performance and subjective ratings by restoring this tactile-vestibular harmony. This work establishes that handling qualities are fundamentally an Information Theory problem: Level 1 ratings occur when the residue-to-decay gradient is tuned to map the dominant mode's physical 'elbow' onto the human sensory window. This paper establishes that perceptual fidelity is not determined by the closed-loop vehicle residues, but by the force sensitivity residues. By identifying the stick-force channel as the superior conduit for this mapping, VVC allows designers to restore informational integrity in virtual environments across all axes of aircraft dynamics. VVC ensures that the visual channel supports a control strategy where the tactile feedback remains within the human sensory window, preventing the 'wrong reading' that leads to simulator-to-flight rating mismatches.
This paper presents the development, optimization, and flight test validation of a Trajectory Control System (TCS)-based flight control system for a tiltwing unmanned aerial vehicle. The TCS is a configuration-independent middle-loop longitudinal controller for vertical takeoff and landing aircraft and is integrated here with explicit model following inner-loop controllers, inverse propulsor models, and a tiltwing-specific control allocation scheme. The resulting flight control system provides coordinated control across vertical flight mode, hybrid flight mode, transition flight mode, and forward flight mode while relying on a concise feedback set and requiring only airspeed from the air data system. The control laws are obtained using a formal constrained optimization framework and transferred directly from simulation to flight without additional on-site retuning. Flight test results from piloted, semi-autonomous, and fully autonomous operations demonstrate stable and predictable behavior throughout the flight envelope, including tight hover performance, simultaneous climb rate and speed tracking in hybrid flight, and successful departure and arrival transitions at multiple speeds. Selected simulation-versus-flight comparisons further show that the nonlinear model captures the dominant trends in the measured response while also identifying specific aerodynamic and transition regime effects that warrant further refinement. Overall, the results demonstrate that the TCS + EMF architecture provides a practical and effective control solution for tiltwing VTOL aircraft.
This paper investigates the impact of aerodynamic interactions on the whirl flutter boundary of wing-twin-propeller configurations. A coupled wing-pylon-propeller model is developed in the Rotorcraft Comprehensive Analysis System (RCAS), where the wing is modeled using uniform inflow and the propeller wake is modeled using the viscous vortex particle method (VVPM). The study examines the effects of spanwise propeller placement and rotation direction by first analyzing a single-propeller configuration and subsequently extending the analysis to twin-propeller configurations. The analyses are performed for both rigid and flexible wings, with the latter designed such that whirl flutter governs the instability boundary. Results show that spanwise propeller placement strongly influences whirl flutter stability, with outboard locations exhibiting higher flutter speeds. Aerodynamic interactions between the wing and the propeller are found to be generally destabilizing, reducing the whirl flutter speed across all spanwise locations, with the largest reduction observed at outboard positions. In twin-propeller configurations, amplitude modulation is observed due to coupling between closely spaced modes enabled by wing flexibility and aerodynamic interactions. The inboard propeller is found to govern the instability boundary across all spanwise arrangements. Inter-propeller aerodynamic interactions are generally destabilizing, particularly for closely spaced configurations, while contra-rotating propeller configurations exhibit slightly higher flutter speeds due to the reduced impact of interaction effects. These results provide insight into the directional nature of aerodynamic interaction effects on whirl flutter in wing-twin-propeller configurations, highlighting their configuration-dependent behavior and generally modest influence on the whirl flutter boundary.
This paper presents the design, development, and successful demonstration of the first-ever ballistically tube-launched tailsitter unmanned aerial system. The vehicle expands upon the capabilities of existing tube-launched systems by simultaneously integrating the hovering capability of a rotary-wing aircraft with the efficiency and speed of a fixed-wing aircraft. To achieve this, the platform's design incorporates a novel coaxial thrust-vectoring propeller system for control in vertical flight and a unique foldable wing design for ultra-compact storage in the launch tube. The aeromechanics of the foldable wings during deployment are studied through a combination of wind tunnel experiments and flight dynamics model simulations, and the results are used to formulate a methodology for executing the ballistic launch. Simulations are also performed to characterize the robustness of the system against asynchronous deployment of the left and right wings. Experimental data, collected from flying a prototype in vertical, horizontal, and transitioning flight, demonstrate the aircraft’s flying performance. The study ultimately culminates with a demonstration of the prototype being rapidly launched from a tube at 25 m/s (56 mph) and autonomously unfolding, stabilizing, and transitioning into self-powered cruising flight.
Advanced air mobility (AAM) seeks to develop a large-scale transportation system to revolutionize how people live and work, with electric vertical take-off and landing (eVTOL) aircraft serving a central role due to reduced emissions and noise impact. An important aspect for eVTOL aircraft certification is safe urban operations, which require understanding of the response due to aerodynamic disturbances. Experimental data are required to support eVTOL aircraft development with respect to flight dynamics and controllability, as well as design specification development. While flight testing of the full-sized air vehicle will be necessary as part of the certification process, subscale testing offers many advantages with respect to cost and flexibility, in addition to examining operational conditions that one would be reluctant to test in flight at full scale such as emergency conditions. These advantages only may be seen if the underlying scaling principles of flight dynamics / control, aerodynamic interactions, and propulsion-airframe integration are understood. This paper describes initial work towards development of a general subscale testing methodology for eVTOL aircraft flight dynamics and disturbance response characteristics including limited degree of freedom (DOF) and free flight testing. An overview of the initial development work is provided, including discussion of scaling relationships, subscale air vehicle model development, and testing activities focusing on flying qualities and stability / control characteristics.
This study investigates the post-failure flight dynamics of a 1200 lb classical octocopter under single motor inoperative condition using nonlinear time-domain simulations with a baseline feedback controller. A physics based propulsion sizing strategy is developed using IEC duty cycle definitions where continuous requirements are derived from nominal hover with margin and short time capability is used to accommodate elevated post failure loads. The selected motor satisfies both regimes and enables transient overdrive without excessive weight penalty. Simulation results in hover and forward flight at the best range speed showing that the vehicle can recover from any single motor failure and retrim using inherent redundancy without fault identification. However, recovery involves significant transient attitude excursions and altitude loss, and requires substantial increases in motor power, with multiple motors exceeding S1 power limits. Post-failure maneuver simulations indicate retained controllability with some degradation and increased coupling. These simulations demonstrate that the proposed motor sizing enables necessary operation post-failure while avoiding unnecessary oversizing.
An advanced coupling framework was leveraged to assemble analytic sensitivities of lifting line theory aerodynamic loads with respect to externally-defined blade geometry parameters for optimization of main rotor performance of conventional helicopter configurations. Three vehicle weights and two flat-plate-equivalent drag configurations were examined across the flight envelope from hover to an advance ratio of 0.3. Two types of twist controls were investigated: quasi-static and fully active. Power savings were strongly correlated to the forward flight to hover power, ranging between 1.5 and 3.5% for quasi-static geometries and 2.0 and 4.5% for fully active controls when the installed power is twice of that required in hover. Blade twists optimized at higher power ratios were observed to favor high shaft tilt angles. Optimal twist deformation relative to hover-optimized designs is nonlinear across the blade span. Minimal penalties to aerodynamic vibrations were incurred through the use of either quasi-static or fully active twist controls as measured with a vibration intrusion index.
A new modeling and simulation approach for uncrewed aerial vehicles (UAVs) for shipboard operations, as well as other complex airwake environments, is introduced. The approach couples an aerodynamic modeling framework, expanded with a closed-loop flight controller, with a flight dynamics solver to assess UAV responses in highly complex airwakes. Using high-fidelity wall-modeled large eddy simulation airwakes, the approach has been evaluated for two traditional naval approaches on a generic destroyer. As part of this effort, a new metric to ensure accurate flight modeling is introduced, along with proposed UAV safety mapping. These predictions can support mission planning by evaluating unique prescribed approaches for nontraditional small UAV, and can aid in early controller design by highlighting specific failure modes based on the airwake environment.
This paper presents an adaptive model predictive control (MPC) framework for nonlinear urban air mobility (UAM) vehicles operating across the full flight envelope. The proposed approach leverages a linear parameter-varying (LPV) representation to update the predictive model online, enabling accurate capture of strongly nonlinear and time-varying dynamics associated with distributed electric propulsion (DEP) eVTOL aircraft. To systematically address the highdimensional and coupled nature of MPC tuning, a multi-objective evolutionary optimization strategy based on NSGAII is employed, incorporating proper normalization of states and control inputs to ensure balanced weighting and meaningful exploration of the design space. The resulting controller explicitly accounts for actuator constraints and enables reconfigurable control allocation for fault-tolerant operation. The framework is evaluated in nonlinear simulations using NASA's Generic Urban Air Mobility (GUAM) model and benchmarked against a robust servomechanism linear quadratic regulator (RSLQR). Results demonstrate that the proposed adaptive MPC achieves improved trajectory tracking and enhanced robustness under both nominal conditions and actuator degradation scenarios, including partial motor failure, while maintaining constraint satisfaction throughout all flight regimes.
Electric Vertical Take-Off and Landing (eVTOL) aircraft are poised to transform urban and regional mobility by offering zero-emission, congestion-free transportation. As regulatory frameworks evolve and advanced air mobility (AAM) gains traction, manufacturers are exploring propulsion strategies that improve range, power delivery, and overall system efficiency. A key challenge in eVTOL development is balancing range with payload capacity. While larger battery packs can extend range, they also increase system weight, reduce payload, and prolong charging times, limiting operational flexibility and turnaround time. Hydrogen fuel cells, supported by liquid hydrogen (LH₂) present a promising alternative for eVTOL propulsion. This study proposes a methodology for optimizing fuel cell propulsion systems tailored to eVTOL applications. A multi-physics modeling framework for eVTOL flight dynamics and propulsion system was developed, representing the target eVTOL configuration. For a defined flight path including vertical takeoff, hover, cruise, and landing, a Genetic Algorithm (GA) based optimization was conducted on propulsion system. The algorithm down-selected battery size, fuel cell stack specifications, and hydrogen tank capacity to meet mission requirements while minimizing propulsion system weight. The modeling framework was also used to evaluate trade-offs between payload and performance as functions of component sizing, battery chemistry and energy distribution strategy.
Unmanned aerial vehicle (UAV) primary structures require high specific strength and stiffness, traditionally necessitating expensive carbon fiber composites. This study evaluates simulation-driven, additively manufactured polymer alternatives fabricated from PLA and computationally optimized via macroscopic Topology Optimization (TO), mesoscopic Variable-Thickness Lattices (VTL), and uniform Triply Periodic Minimal Surfaces (TPMS). Evaluations were conducted under a superimposed, multi-axial flight envelope. Physical testing demonstrated that VTL architectures maximized the Structural Efficiency Index (SEI) by pushing mass to the extreme geometric fibers and increasing global flexural rigidity. In contrast, mass-constrained TO yielded misleading specific strength due to volumetric starvation and elevated compliance, while the uniform TPMS baseline exhibited favorable specific stiffness but lacked targeted root robustness, resulting in reduced specific strength. Off-axis testing further showed that Diamond lattices dominated vertical bending and inverted impulse loading, whereas Octet and Kelvin geometries more efficiently resolved transverse shear. Experimental data identified a performance-optimized efficiency asymptote in the 73-79 g VTL specimens, which achieved a 19-24% mass reduction relative to a 97 g carbon fiber baseline. To assess assembled-vehicle relevance, the selected fully 3D-printed replacement arms were installed on the baseline quadcopter and subjected to nine dynamic ground tests comprising staircase and cyclic propulsive loading under freestream conditions of 0, 10, and 20 knots. The optimized arms completed the full test matrix without fracture, mount failure, screw loosening, visible yielding, or permanent deformation, demonstrating structural viability in a realistic multi-part UAV assembly without carbon fiber reinforcement.
This study investigated the feasibility of using Deep Reinforcement Learning (DRL) for aeroelastic stability control of a Tiltrotor Aeroelastic Stability Testbed (TRAST) model. The DRL controllers use rotor swashplate inputs to minimize oscillatory wing root bending moments of the tilt rotor model. First, three DRL-based agents including Deep Deterministic Policy Gradient (DDPG), Twin Delayed Deep Deterministic Policy Gradient (TD3), and Soft Actor-Critic (SAC) were investigated to control the aeroelastic stability of the TRAST model throughout a wide range of airspeed including where the whirl flutter occurs. All three agents demonstrated the capability of stability augmentation while the SAC agent demon-strated the most robust performance. Next, the effectiveness of the SAC agent was studied further by training the SAC agent at a certain airspeed and applying the trained agent through the TRAST whirl flutter conditions. Finally, additional tuning of the SAC agent was performed to improve performance further through a hyperparameter optimization framework called Optuna.
A method for evaluation of control derivatives is introduced for the purpose of rapid design evaluation of an electric, fixed-pitch multirotor aircraft during the conceptual pre-design phase. This explicit linearization methodology allows rapid co-design of the vehicle configuration and control allocation using the pseudo-inverse method. A multi-objective design analysis is conducted for a 12 rotor lift + cruise eVTOL configuration subject to hover power requirements, controllability, and tolerance to failure conditions. Generalizable design guidelines are found and presented for the cant and rotor spin direction of the lift + cruise aircraft. The benefits shown include the addition of direct lateral force control derivative, a major increase in yaw control derivative, and reconfiguration to accommodate any Two Engine Inoperative failure conditions. These are achieved through mixing anhedral and dihedral rotor cant within each quadrant of the wing, setting the spin direction so the component of thrust in the yaw axis is additive with the motor reaction torque, and having larger cant angles inboard and closer to the center of gravity. Additional analyses of stability characteristics are performed using DynaPyVTOL, a medium-fidelity flight dynamics analysis tool for any configuration.
The paper discusses the control law synthesis problem for the Advanced Tiltrotor Aircraft (ATA), specifically considering the lateral-directional dynamics in Airplane mode. The aircraft under investigation is a new tiltrotor concept developed by Leonardo Helicopters. The ATA aims at positioning itself as a new evolution of the tiltrotor concept, following the experience matured with existing aircraft configurations, as the AW609 and the NGCTR. During the concept design phase of a complex vehicle such as a tiltrotor, it is neither possible nor encouraged to rely on high-fidelity flight mechanics models (e.g. multibody), due to their computational cost and the large number of design parameters required, which are not typically known with sufficient accuracy in this project stage. Therefore, technology validation activities in the field of flight dynamics and control should rely on lean and flexible simulation tools, based on reduced set of design parameters but yet capable of predicting most important dynamic characteristics in the frequency range of interest for flight controls. The paper first addresses the inclusion of rotor gimbal states in the original ATA 6 dof flight dynamics code, which based on Tiltrotor experience are considered crucial for a proper identification of dominant aircraft dynamics in high inflow conditions, i.e. Airplane mode. The outcomes of lateral-directional model trimming and linearization procedures are then assessed, and a novel basic lateral-directional Command Augmentation System scheme designed and tested, in both linear and non-linear framework, against prescribed MIL-STD-1797B performance and dynamic stability requirements and SAE-AS-94900 robust stability criteria. The results show that proposed control strategy is a viable approach for further investigations and the adopted flight physics modelling complexity adequate for the addressed control problem.
This paper presents the integration and use of state-space free-vortex wake models within closed-loop rotorcraft flight dynamics simulations. The free-vortex wake models are formulated in state-variable form, such that they constitute a system of nonlinear, time-varying ordinary differential equations in first-order form that augment the baseline rigid-body and rotor dynamics. The wake models considered include a tip-vortex-only formulation, as well as a formulation combining a vortex-lattice near wake with a tip-vortex representation of the far wake. Following trimming, linearization, and model-order reduction of the flight dynamics at discrete increments in flight speed, Dynamic Inversion (DI) flight control laws are synthesized to enable automatic transition from hover to cruise flight. Two-way-coupled losed-loop simulations are then performed for a generic utility helicopter representative of an H-60 in transition from hover to forward flight using three inflow models: (i) Pitt-Peters, (ii) tip-vortex-only wake, and (iii) vortex-lattice near wake with tip-vortex far wake. These simulations are compared both for validation purposes and to assess whether the different wake representations lead to significant differences in the predicted closed-loop response.
This paper presents the development flight test campaign of autopilot Upper Modes for T-625 Gökbey helicopter. The primary objective of the test campaign is to evaluate the newly developed Upper Modes in the frequency and time domain across the operational flight envelope. For quantification of performance and stability, various metrics are selected from the literature. Flight tests are designed to extract the metrics from time domain data and tests are conducted. Initial flight tests revealed discrepancies between theoretical design models and actual aircraft dynamics, requiring iterative control law gain optimizations. Furthermore, combined mode engagements required targeted simultaneous tuning of different modes to maintain stability margins in combined engagement. By integrating quantitative data analysis with qualitative pilot feedback, engagement logic and control parameters were successfully refined.
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.
This paper presents the integration of a state-space viscous vortex particle method (SS-VVPM) into rotorcraft flight dynamics simulations. More specifically, the near wake is modeled using a near-wake vortex-lattice method, while the far wake is represented by viscous vortex particles. The VVPM equations are formulated as a nonlinear, time-varying system in first-order form, allowing the wake dynamics to augment those of the rigid body and rotors. This formulation enables the coupled vehicle, rotor, and wake dynamics to be trimmed and linearized for use in stability analysis and flight control design. Different strategies for coupling the near wake to the particle wake are investigated, and a tip-edge shedding approach is identified as the best compromise between computational cost and accuracy. Verification is carried out for a generic utility-helicopter rotor in hover and forward flight through comparisons with previously validated free-vortex wake methods. Results show good agreement in wake geometry, thrust prediction, and tip-vortex trajectories, indicating that the particle-based formulation retains the dominant wake physics of the legacy model. Additional simulations for generic helicopter, tiltrotor, and electric vertical takeoff and landing (eVTOL) configurations over a range of flight speeds further demonstrate that the method reproduces the expected changes in wake structure from hover to forward flight.
This paper presents the development of a ballistically launchable, 360 g coaxial Micro Air Vehicle (MAV) utilizing a two-axis gimbal thrust-vectoring system for pitch and roll control. A combined experimental and analytical framework is employed to characterize aircraft performance across the full ballistic profile, from the initial projectile phase to the transition to hover. Wind tunnel experiments are first used to quantify passive stability during the projectile phase. Subsequently, a nonlinear six-degree-of-freedom (6-DOF) flight dynamics model is developed by synthesizing the mechanical and aerodynamic models of the individual aircraft subsystems. These subsystem aerodynamic models are obtained by combining experimental lookup tables with analytical models. The flight dynamics model is then leveraged to analyze the transition phase, characterize the effectiveness of the thrust-vectoring mechanism, and establish performance guarantees across the ballistic flight profile. A key finding of this work is that the effectiveness of the two-axis gimbal thrust-vectoring mechanism is fundamentally limited by the axial velocity experienced by the propellers at the point of transition. This operational limit is quantified for the coaxial MAV, and the vehicle's ability to successfully transition to hover during a ballistic launch is validated through a single-degree-of-freedom (1-DOF) experimental testing.
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.
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.
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.
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.
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.
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.
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.
The wing-in-ground effect (WIG) vehicle represents a significant advancement in aerodynamics and vehicle design, leveraging the ground effect phenomenon to enhance lift and reduce drag when flying close to the surface. This unique capability allows WIG vehicles to achieve higher payloads, longer range, and greater fuel efficiency compared to traditional aircraft, making them an attractive option for modern military and global disaster response applications. Wing-in-Ground Effect Vehicles: From Modern Military and Commercial Development to Global Disaster Response discusses future disaster response, logistics, and military applications for WIG vehicles, including the ongoing development of aerospace and transportation technology. Relavant advancements in materials and propulsion systems holds promise for further enhancing WIG performance and operational range. Additionally, cost-effective and powerful flight computers with various types of mission-enabling sensor suites from the advanced air mobility and drone industry further enable the development of modern, safe, and practical WIG platforms. Click here to access the full SAE EDGETM Research Report portfolio.
This paper presents the results of an ongoing correlation study performed using three different comprehensive rotorcraft codes and data obtained from the Advanced Testbed for TILtrotor Aeroelastics (ATTILA) tiltrotor whirl flutter wind tunnel test campaign. The ATTILA testbed consists of a 1:5 scale semi-span wing with a powered, tip-mounted proprotor reflecting the proprietary design of the Next Generation Civil TiltRotor (NGCTR). Experimental dynamic characterization of the testbed has revealed non-negligible structural nonlinearities. Post-test efforts have focused on refining the damping trends extracted from the test data, and correlating the experimental results with numerical predictions. The objective of this paper is to assess the modelling fidelity required and afforded by modern comprehensive aeromechanics codes to predict tiltrotor whirl flutter instability given an industry-representative design that exhibits structural nonlinearities. Baseline numerical flutter models fail to predict some of the observed experimental damping behaviour, but the inclusion of higher fidelity aerodynamics and exploratory friction models improves prediction accuracy. Ongoing modelling and dynamic characterization efforts aim to further clarify the mechanisms influencing the whirl flutter stability of the ATTILA testbed and enhance the predictive capabilities of the numerical methods employed.
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.
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.
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