Flying-wing and tailless UAVs can integrate the lifting surface, structure, payload volume, propulsion, and control surfaces into a compact airframe with fewer nonlifting components. Their design places special demands on the spanwise loading because the wing must also provide trim, structural efficiency, and control without a conventional empennage. A unified closed-form framework was developed that relates the prescribed spanwise circulation distribution to the induced velocity, local and integrated induced drag, wing-root bending moment, span-resizing trade, and geometric twist required to realize the loading. Closed-form twist relations were derived for both constant-chord and linearly tapered wings. For fixed lift and span, elliptical loading remains the minimum-induced-drag solution. The Prandtl-type loading has higher integrated induced drag but produces outboard upwash and local negative induced drag, allowing differential outer-wing lift to contribute to proverse yaw. The same inboard lift redistribution reduces root bending moment, allowing span or aspect ratio to be increased at the same idealized root bending moment.
In-flight, real-time ice detection enhances the pilot’s airframe icing awareness during aircraft and unmanned aerial vehicle operations in adverse weather. The objective of this work is to develop logistic regression–based binary and multiclass classifiers for ice class predictions to determine the onset of glaze ice horns on an airfoil using the degradation of aerodynamic performance during icing at mid- and low-Reynolds-number conditions. Multiclass classifiers divided the glaze ice horns class into incipient and developed. Cross-validation results yielded validation accuracies of 97% and 71% for ice horns onset detection accuracy. The high validation accuracy indicates that the instantaneous ice class can be inferred from the instantaneous aerodynamic performance. The glaze ice horns onset accuracy of the models was improved by 9% by overfitting the streamlined rime to glaze ice horns transition and by using the aerodynamic performance time evolution as input. It was observed that a larger training batch reduced the glaze ice horns onset accuracy when using overfitting strategies. The trained models were used in virtual test cases showing that the improved multiclass classifier can anticipate the glaze ice horns onset by sensing a variance increment of the glaze ice horns probability.
Effects of acceleration and deceleration on aerodynamic characteristics at levels typical for commercial aircraft have not yet been investigated. This study examines these effects by solving the two-dimensional Euler equations on moving Cartesian meshes that follow a flying body. Computational accuracy is validated by comparing computed standoff distances of a bow shock wave ahead of a decelerating sphere with results from previous experiments and simulations. Subsequently, unsteady flow simulations are performed around a NACA 0012 airfoil with a chord length of 20 m under standard atmospheric conditions at an altitude of 10 km. A flight Mach number varies between 0.5 and 1.2, with acceleration and deceleration rates of [Formula: see text], [Formula: see text], [Formula: see text], and [Formula: see text]. The results reveal distinct flow behaviors among cruising, accelerated, and decelerated flight. In accelerated flight, the aerodynamic coefficients exhibit trends similar to those in cruising flight. Conversely, in decelerated flight, they differ significantly in transonic flow regimes due to the persistence of the rear shock wave, followed by its sudden upstream movement. This leads to a hysteresis phenomenon between accelerated and decelerated flight. This hysteresis is amplified with increasing acceleration and deceleration rates and is largely independent of the angle of attack.
The research and development of natural laminar flow (NLF) wing designs can minimize airframe drag. This investigation extends to the design of NLF vertical tails for a reference subsonic civil aircraft under cruise conditions. Although parameters unrelated to design are treated as constant, practical aircraft development requires performance optimization across all airframe components rather than isolated elements. This study examines how the presence and positioning of a horizontal tail affect the turbulent transition and aerodynamic properties of a NLF vertical tail. Adjusting the presence and location of the horizontal tail alters the flowfield around the vertical tail and modifies the spatial pressure distribution, which corresponds to changes in flow acceleration and deceleration. Compared to the baseline configuration, positioning the horizontal tail upstream and vertically higher locally delays transition. The laminar flow extent ranges from approximately 45 to 48% across the configurations, while the transition characteristics are notably affected. The drag coefficient decreases, whereas the lift and moment coefficients and the lift-to-drag ratio also decrease. These results demonstrate that modifying the relative positioning of adjacent components can alter pressure distributions and transition characteristics without changing the geometry of the target component.
This paper presents a process for a consistent coupling of aeropropulsive and engine performance analyses for conceptual design of low-boom supersonic aircraft. An inviscid computational fluid dynamics (CFD) solver is used for aeropropulsive analysis. A variable cycle engine (VCE) model and a low-boom supersonic aircraft are selected to demonstrate the coupling process. The main technical contribution is a block coordinate optimization method to generate an axisymmetric, external-compression, supersonic inlet for a cruise Mach of 1.8, with a total pressure recovery of 97.2%, that has the same aerodynamic interface plane shape and approximately the same mass flow rate as the VCE model at start of cruise (SOC). A numerical optimization of the nozzle shape is also performed to generate a CFD engine that emulates the VCE at SOC, achieving a consistent coupling of aeropropulsive and engine performance analyses for conceptual low-boom design. Off-body pressures and undertrack sonic boom ground signature of the aircraft with two CFD engines are compared to those of the same aircraft with two flow-through nacelles. The documented results illustrate that the aeropropulsive simulation has a significant influence on near-field pressure calculation, but less so on the overall shape of the undertrack sonic boom ground signature.
This study presents computational analyses of coaxial rotor hub flows and validation against experimental data obtained from the rotor hub flow prediction workshops. Experiments were conducted in a 12-in.-diam water tunnel at the Pennsylvania State University Applied Research Laboratory, employing tomographic particle image velocimetry and precise hub drag measurements. The computational fluid dynamics framework which uses an unstructured/Cartesian multimesh paradigm and hybrid Reynolds-averaged Navier–Stokes/large-eddy simulation modeling was applied to replicate and analyze hub flows. Two different hub configurations, consisting of counterrotating hubs, fairings, and shafts, were simulated and compared to experimental results, with a focus on hub drag, wake velocity fields, and turbulence quantities. Results demonstrated that the computational frameworks effectively captured key flow physics, although some discrepancies in drag harmonics, wake velocity, and turbulence intensity magnitudes were observed. Additionally, the study highlighted the impact of rotor hub geometry and installation of a sail fairing on drag and wake structures. These findings contribute to improving computational predictions, essential for designing high-speed rotor hub configurations.
Drawing on the cone-derived method, this study established a waverider design approach with a given leading edge from axisymmetric flows. This approach takes the given leading edge as input to generate a rotated shock surface and a corresponding two-dimensional shock profile. Then, an axisymmetric flow behind the shock profile is simulated using the inverse method of characteristics (iMoC). Accordingly, the waveriders with the given leading edge are generated, wherein streamline-tracing efficiency is improved through quadtree partitioning. Taking a leading edge with wing dihedral expressed by a B-spline as the baseline, the waveriders were generated, and the effect of the rotational axes was analyzed. Results show that the leading edge of a waverider can be solely determined by the shock profile of an axisymmetric flow. Waveriders designed using axisymmetric flows exhibit high [Formula: see text] ratios at hypersonic speeds, demonstrating excellent waveriding performance. The axisymmetric flows with different rotational axes made the waveriders with different thicknesses, showing a strong correlation between the [Formula: see text] ratio and the volume ratio. When the offset distance of the rotational axis increased, the longitudinal stability of the corresponding waveriders decreased. While the waverider with the given leading edge improves its design flexibility, current deficiencies of the iMoC and axisymmetric flow properties impose limitations onvolume expansion and precise control of leading edges.
Feather-like ice on the inlet spinner of an aeroengine endangers flight, yet its formation mechanism was unknown. This study systematically investigates the formation mechanism of ice accretion on a conical spinner using icing-wind-tunnel experiments and 3D scanning technologies. Experiments revealed two distinct ice morphologies on the spinner: forward-growing nail-like feathers and backward-growing needle-like feathers. The transition between these morphologies was found to depend critically on rotational speed, liquid water content (LWC), incoming flow velocity, and temperature. Specifically, higher rotational speeds, elevated LWC, increased flow velocities, and warmer temperatures promoted the shift from forward nail-like to backward needle-like ice structures. Three-dimensional characterization of ice profiles identified two key angles (alpha and phi) governing growth directionality, with alpha defining windward or leeward inclination and phi representing lateral broadening perpendicular to streamlines. Under rotation, centrifugal forces and shear stresses redistributed unfrozen water films on ice particles, driving leeward growth of needle-like feathers. When the spinner is stationary, the merging of adjacent ice elements along the phi direction generates broader structures, forming ice structures analogous to scallop ice on swept wings. These findings provide the first 3D experimental evidence of dynamic ice-feather evolution on rotating spinners, resolving longstanding gaps in understanding spinner-specific icing mechanics.
The near-ground flight of new-generation aerospace vehicles-such as horizontally launched electromagnetic systems and sea-skimming missiles-requires a comprehensive understanding of high-speed ground effect (GE) aerodynamics. However, transonic and supersonic GE are far less studied than the subsonic regime. This work systematically examines GE aerodynamics of a representative airfoil across subsonic, transonic and supersonic regimes and a wide range of flight heights by numerically solving the compressible Reynolds-Averaged Navier-Stokes equations. The evolution of flow patterns and aerodynamic characteristics with Mach number in GE is provided to complement classical textbooks focusing on unbounded flows. Relative to subsonic GE, aerodynamic force variations in the transonic and supersonic regimes are substantially larger. The conventional critical height criterion of h/c approximate to 1 becomes invalid at higher Mach numbers, where significant GE persists even at elevated heights. As the height decreases, the lower critical Mach number increases slightly, whereas the upper critical Mach number rises sharply. Across all regimes, the total lift is primarily governed by pressure changes on the lower surface, although the dominant mechanisms differ: in the subsonic regime, the blocking and camber effects prevail, whereas in the transonic and supersonic regimes, the complex shock-ground interactions are the principal contributors.
The AIAA workshop series supports the computational fluid dynamics (CFD) community through impartial evaluation of publicly available test case submissions to establish and improve the state-of-the-art in applied CFD. The Fifth AIAA CFD High Lift Prediction Workshop (HLPW-5) is organized to assess the numerical prediction and physical modeling capabilities of CFD technology for swept wings in landing and takeoff high-lift configurations. Multiple Common Research Model High-Lift variants are assessed, and the Mesh Adaptation Technical Focus Group submissions are summarized. Test Case 1 focuses on turbulence model verification. Variation between methods is low for a fuselage with a single-element wing, which allowed for convincing verification of two Spalart-Allmaras turbulence model variants. Test Case 2 provides an incremental high-lift component buildup to isolate the influence of geometry. Variation between methods increases slightly with a deployed leading-edge slat and increases dramatically with the addition of trailing-edge flaps. Comparisons to experiment and between submissions are favorable for the configuration without trailing-edge flaps at low and moderate angles of attack, which roughly defines the validation and validity threshold of existing methods using steady Reynolds-averaged Navier-Stokes (RANS) turbulence models. Documenting this validity threshold sets the stage for developing RANS modeling improvements.
Research on reliability-based design optimization (RBDO) that considers solely input variable uncertainty is well-established. However, studies on RBDO problems that involve both random input variables and their uncertain distribution parameters remain notably scarce. This study constructs an augmented-RBDO (A-RBDO) framework governed by target augmented failure probability (AFP) constraints under double-layer uncertainty. To enable its efficient solution, this study proposes an efficient sequential decoupling method for solving the A-RBDO model, utilizing augmented quantiles of the performance function. The proposed method first equivalently transforms the AFP constraint into an augmented quantile constraint. This transformation converts the A-RBDO problem into a sequential process involving augmented quantile estimation and deterministic optimization. Subsequently, for critical augmented quantile estimation, this study introduces an efficient strategy combining dimension-reduction integration based on a mixed-degree cubature formula with the bisection method, significantly reducing the expense associated with augmented quantile evaluation. The synergistic integration of this sequential decoupling strategy, the mixed-degree cubature formula, and the bisection method dramatically enhances efficiency in solving the A-RBDO model. The efficiency and robustness of the proposed method are comprehensively demonstrated through three examples.
This paper presents a summary of the computational flutter results associated with the AIAA Third Aeroelastic Prediction Workshop, High Angle Working Group. The computational results are compared against the experimental data collected during the Pitch and Plunge Apparatus Benchmark Supercritical Wing test campaign conducted in the Transonic Dynamics Tunnel at NASA Langley Research Center in 1993. During that test, several flutter points were identified at transonic conditions. One of these points, specifically near Mach 0.8 and a 5 deg angle of attack, became a focal point of the computational challenge within the working group. Various-fidelity time-domain, reduced-order model, and linearized frequency-domain methods were used by seven participating teams, and a description of each team’s software and methods is included. While there are encouraging trends in the computational results, the range of the predicted flutter dynamic pressure is still quite large due to the stall flutter mechanism. Comparisons are also complicated by the potential existence of a strong limit-cycle oscillation: the strength of the aeroelastic damping may depend strongly upon the size and character of the applied perturbation, but this perturbation had not been specified as a fixed parameter for workshop participants.
Among the most versatile lighter-than-air platforms, unmanned airships are currently being designed mostly for either low-altitude missions for close-distance surveillance, in competition with multicopter drones, or for high-altitude missions in the stratospheric layer, thus ideally complementing the role of space satellites. Correspondingly, algorithms to automatically compute global values such as the volume and mass of an airship for a desired mission performance and for assumed technologies of the components (such as the materials employed for the envelope or solar cells) have been experimented with and are documented in the literature. Building on this base, this research proposes a method where not just the parameters most typical of preliminary design are solved, but a unified automatic approach is employed to take into account requirements on static balance as well as dynamic performance in the form of characteristic time and damping of some of the eigenmodes. In this way, the outcome of the automatic sizing procedure may account not just for the requirements of the mission profile, but also potentially for static balance and for a desired level of flying qualities.
The ability to control the flow in the engines of aircraft in crossflow conditions is especially important for future aircraft equipped with ultra-high-bypass-ratio engines. This challenge is addressed computationally by employing the traverse actuation method to enhance flow quality in the inlets of turbofan engines and electric fans. This unsteady actuation is based on an array of discrete jets blowing in the general streamwise direction toward the fan face. The jets move in the circumferential direction. The continuous actuation produces small, fast-moving jet packets that travel laterally in a periodic manner, resembling the sideways motion of a crab. This spatial sweeping motion along the inlet surface provides extensive area coverage and is very effective in reducing flow separation in a quasi-steady manner. Special time-varying boundary conditions are used to simulate the traverse actuation. The efficiency of this flow control approach is demonstrated for an aircraft engine cowl in crossflow, where flow separation that originally develops on the windward side is effectively reduced or outright eliminated using a small level of actuation input. The study prioritizes low actuation input, which is essential for practical system integration. Actuator devices that provide the desired flow control patterns are also described.
Tilt-rotor propulsion system design requires a multidisciplinary approach to tackle important challenges and competing tradeoffs between disciplines. This paper models rotor aerodynamics, blade structures, vehicle drag, electric propulsion, and tonal/broadband acoustics for a tilt-rotor, electric vertical takeoff and landing aircraft using low-to-mid fidelity tools. The authors use gradient-based design optimization with automatic differentiation and parameter sensitivity analyses to explore the design space and complex tradeoffs of tilt-rotor distributed electric propulsion systems, exploring effects of variations in payload/empty weight, battery specific energy, and blade tip speed. This framework models multiple operating points with a mission-focused objective to account for the effects of both hover and cruise conditions on the overall system performance. Additionally, we develop a Pareto front between range and noise and observe that, for the same noise output, modeling tonal and broadband noise increases range by 3.1% when compared to using a Mach tip speed surrogate acoustics model.
An aeroelastic dataset generation method for aircraft based on coupled computational fluid dynamics-computational structural mechanics (CFD-CSM) simulations and a Chimera approach for control surfaces is presented. The procedures are embedded in the multidisciplinary simulation environment SimServer. The TAU code is employed to solve the Reynolds-averaged Navier-Stokes equations. Structural displacements are computed with a modal solver. The dataset production tool is applied to a reconstruction of the Northrop F-5A aircraft, representing aerodynamic characteristics, including static aeroelastic effects. Control surface deflections of the leading-edge flaps, trailing-edge flaps, ailerons, and horizontal stabilizers are considered. The effects of control surface deflections on the flowfield as well as on the force and moment coefficients are analyzed. The influence of structural elasticity on the aircraft configuration is evaluated by comparing the aerodynamic coefficients of the aeroelastic simulations to those of the purely aerodynamic simulations, for which the aircraft is assumed to be fully rigid. In addition, variable dynamic pressure and variable inertia loads at several control surface deflection combinations are investigated. The results show that the structural elasticity of the aircraft significantly influences the aerodynamic force and moment coefficient characteristics for large regions of the flight envelope.
Drag minimization of aerodynamic shapes in transonic flow can lead to shock-free solutions with poor off-design performance. The work here explores whether optimizing range, augmented by the addition of the operating point as a design variable, is an appropriate objective for inviscid transonic wing optimization. Gradient-based optimizations are performed, showing that the range formulation allows the optimizer to increase drag divergence to maximize speed, which in turn allows a lower lift coefficient to minimize induced drag. The result is a shocked solution, with the shock moving further aft on the wing for higher weight, which can be managed via a pitching moment constraint. Range optimizations lead to higher optimum Mach numbers, which raise drag divergence. Similar behavior is seen with multipoint optimization with a high-speed point included, though the range optimizations have better global off-design performance (via a larger region of high performance around the optimum performance point), better local off-design performance (via lower sensitivity in range to changes in speed and lift), and are further from drag divergence.
Aircraft Maneuver Load Alleviation (MLA) is a multidisciplinary problem involving the coupled physics of aerodynamics, flight dynamics, and structural analysis. This work presents an analytical model-based investigation of MLA through movable control surfaces for various wing platforms, aiming to understand the underlying physics, identify key influencing parameters, and inform the formulation of associated multidisciplinary optimization problems. The investigation starts with a rigid wing and a single-point aileron, and incrementally increases modeling complexity by introducing finite aileron span, aircraft trim conditions, alternative lift distributions (e.g., Prandtl’s Bell-Shaped) and wing flexibility to further improve the fidelity. Key findings reveal that optimal aileron placement for load alleviation is highly sensitive to trim effects, shifting from approximately 71% span for a rigid wing to 82% span when trim is considered, and further to 64.33% span for a trimmed wing with a Bell-Shaped lift distribution. Trim effects are identified as the dominant factor due to compensatory pitch angle changes, while wing planform (lift distribution) also has a significant influence. In contrast, the aileron span and wing flexibility (for straight wings) exhibit comparatively minor impacts on optimal placement. Furthermore, the analysis demonstrates that while flaps can also achieve load alleviation with optimal placement near the wing root, the optimal aileron placement is independent of flap usage. These analytical insights provide foundational understanding, crucial for guiding preliminary aircraft design and streamlining subsequent computationally expensive high-fidelity Multidisciplinary Design Optimization (MDO) efforts.
This paper presents a design-oriented investigation of a commercial transport aircraft equipped with semi-aeroelastic hinged (SAH) wingtips. A multibody flight-dynamics formulation is developed to model articulated wings with hinge degrees of freedom, addressing a key limitation in previous SAH analyses. The framework is coupled with a hybrid aerodynamic approach combining experimental baseline data with CFD-validated vortex-lattice method predictions, together with an improved panelization strategy for hinge axes not aligned with the freestream. Two SAH concepts are evaluated: a locked hinge acting as a rigid span extension and a continuously free hinge that passively rotates under aerodynamic loading. For a 10% span increase, the locked hinge provides only a marginal efficiency improvement (0.2% lower required engine power). For a 20% span increase, the locked hinge yields a larger benefit (1.5%). Simulations with sudden angle-of-attack changes show reduced wingtip overshoot for the free hinge, indicating passive load alleviation capability. The proposed framework enables quantitative design-space exploration of SAH wingtip concepts.
