Steady Level Flight Research Articles

Air and Structural Loads Analysis of a 5-Ton Class Rotorcraft in a Pull-Up Maneuver Using CFD/CSD Coupled Approach

The air and structural loads of a 5-ton class light helicopter (LH) rotor in a 2.24 g pull-up maneuver are investigated using a coupling between the computational structural dynamics (CSD) and computational fluid dynamics (CFD) methods. The LH rotor is characterized by a five-bladed system with elastomeric bearings and inter-bladed dampers. The periodic trim solution along with the converged CFD/CSD delta airloads obtained in steady-level flight (advance ratio of 0.287) are used to perform the transient CSD maneuver analysis. The resulting vehicle attitude angles and velocity profiles of the aircraft are then prescribed in the quasi-static (QS) CFD maneuver analysis. It is demonstrated that the present QS approach provides an effective means for the maneuver loads’ analysis. The important flow behaviors such as BVI (blade–vortex interaction)-induced oscillations and the negative pitching moment peaks met in maneuver flight are captured nicely with the proposed method. Either the vortex trajectories or the surface pressure distributions are examined to identify the sources of the oscillations. A loose CFD/CSD coupling (LC) is used to predict the blade elastic motions, structural moments, and pitch link loads at the specified maneuver revolution of the rotor and also to correlate these with the transient CSD-based predictions. A reasonable correlation is obtained. The LC results show more pronounced 5P (five per revolution) oscillations on the structural response than those of the CSD-based methods.

Simulation of tilt-rotor UAV flight dynamics in horizontal flight

Tilt-rotor unmanned aerial vehicles (UAV) have gained significant importance in the aeronautical industry due to their ability to transition between vertical and horizontal flight. One of the important steps in the development of such UAVs is to assess their performance and stability characteristics. Simulation of flight dynamics is an essential tool that enables the designers to test and optimize the flight characteristics of these UAVs. This paper presents a complete procedure for developing a 6-degree-of-freedom flight dynamics model of a tilt-rotor UAV in fixed-wing mode using a physics-based modeling approach. All necessary sub-models are developed and integrated into the Simulink simulation environment to allow for predicting the UAV’s dynamic response. The developed model is verified for correct operation by simulating horizontal steady-level flight and exciting its natural longitudinal modes.

Solid-State Electroaerodynamic Aircraft Design Using Signomial Programming

Electroaerodynamics (EAD) is a form of airbreathing electric propulsion that uses high voltages to produce and accelerate ions, generating thrust without any moving parts. This method of solid-state propulsion is nearly silent and produces no direct combustion emissions. This paper describes a program for designing and optimizing fixed-wing solid-state aircraft propelled solely using EAD. Signomial programming (SP), an emerging method of efficient multidisciplinary design optimization, is employed. The program incorporates performance models for state-of-the-art EAD thrusters, and for custom high-voltage power electronics. It can be used to generate vehicle specifications, or to assess the effect of technological improvements on vehicle performance. The program was used to design and build a flight demonstrator aircraft, based on a goal of achieving steady level flight. In 2017–2018, the demonstrator successfully achieved steady level flight, a first for a fixed-wing electric aircraft with solid-state propulsion. The unproven nature of the EAD propulsion system necessitated a design philosophy centered around minimal technical risk, affecting configuration selection and choice of objective function. A similar design philosophy may prove useful for other design projects with unproven propulsion systems. Finally, endurance values greater than 30 min are achievable with recent improvements in EAD thruster and battery technology.

Aerodynamic Investigation and Simulation Studies on Wing Section of an Unmanned Aerial Vehicle Attached with Solar Plate

The paper investigates the aerodynamic performance and power requirement characteristics of wing sections integrated with high-lift airfoil to support the operation of solar-powered Unmanned Aerial vehicle (UAV). The flight mission is aimed to simulate the operation of solar-powered UAVs under low -speed environment. The research focuses on studying the aerodynamic effect of non-solar UAV wing model and solar UAV wing model for the varying angle of attack. The UAV wing models are tested using a subsonic wind tunnel to validate the aerodynamic characteristics at low-speed condition. The aerodynamic parameters such as coefficient of lift (Cl), coefficient of drag (Cd), coefficient of pressure (Cp), and the total power required to accelerate the solar UAV are studied to maintain steady level flight. The solar UAV and non-solar UAV wing models were subjected to a computational process to examine the pressure and velocity distributions for the aerodynamic performance analysis. Evident results show that the solar cells positioned at the flow separation region of the UAV wing model produces an aerodynamic efficiency rate of 5.45% and required 37.13W of minimum power compared to non-solar UAV at the Reynolds number of 9.8 106.

Numerical Analysis of Glauert Inflow Formula for Single-Rotor Helicopter in Steady-Level Flight below Stall-Flutter Limit

This article addresses the numerical computation problem of induced inflow ratio based on the helicopter momentum theory in forward flight. The Glauert inflow formula (equation) is a nonlinear equation usually solved by the Newton–Raphson method in a relatively small number of iterations. However, many high-order convergence multipoint iterative methods have been developed over the last decade. The study examines several selected methods in terms of finding ones that provide a solution in only one iteration with acceptable accuracy. Furthermore, the influence of initial guesses on the accuracy of the obtained solutions has been investigated. In this regard, the practical range of parameters of the Glauert inflow equation for helicopters in forward flight is roughly determined by simplified modeling of a power and stall-flutter limitation. For these purposes, a basic low-fidelity longitudinal trim model of a single-rotor helicopter in steady-level flight is modified and numerically solved by a symbolic transformation of a system of 20+ nonlinear equations into a single nonlinear equation.

Multi-Domain Based Computational Investigations on Advanced Unmanned Amphibious System for Surveillances in International Marine Borders

The conceptual design, component selection, and deployment experiments of an unmanned amphibious system (US) with a unique Becker in vertical stabilizer based on hydrodynamic research are included in this work. The use of USs is currently expanding significantly, and they are used for fish detection, oceanographic mapping, mining detection, monitoring marine life, and navy purposes. With a maximum forward speed of 30 m/s, the US’s hull is largely built with criteria for identifying and researching marine species. The significant lifetime decline of ocean species drives the deployment of unmanned vehicles for species monitoring from the water’s surface to 300 m below the surface. In addition, the medical team can help the species with health problems using this planned US because they have been identified. The conceptual design and estimated analytical equations encompass the fuselage, Becker rudder, propeller, and other sub-components. The locations of sensors, primarily used to locate mobile marine life, are also considered. A Becker rudder has been imposed to make sharp turns when the US is submerged in water. An advanced hydro propeller produces the propulsion with a 20 cm base diameter. Additionally, a piezoelectric patching-based energy extracting approach is used to the hydro-outside propeller’s surface. As a result, the electrical power generation for different lightweight materials is computed for the performance of US manoeuvrings. With the help of CATIA modelling of the intended USs and ANSYS Fluent hydrodynamic simulations, appropriate high-speed configurations are selected. Various stages of its mission profile, including the US in steady-level flight, the US in climb, and the US over the ocean surface, are subjected to computational simulations. Using an advanced computational technique and previously established experimental correlations, the reliability of these various computational solutions is examined and kept at an appropriate level. This US is highly suggested for marine-based real-time applications due to its acceptable output.

Linear parameter-varying-based transition flight control design for a tilt-rotor aircraft

This paper presents the development of novel transition flight controllers for a class of urban air mobility aircrafts configured with a fixed-wing and six distributed electric rotor assemblies. Only the two tilt-rotors are utilized for thrust vectoring during transition flight from hovering to steady-level flight, while the four lift-rotors are modulated with aerodynamic lift induced by fixed-wing to maintain stable altitude-hold. Three tractable tilt-rotor articulation profiles are proposed by taking into account of various aircraft and hardware constraints. Given a predefined nominal tilting profile, a family of linear models is obtained by linearizing the nonlinear aircraft model at multiple tilt-rotor angular positions along the tilting profile. Using tilt-rotor angular position as a scheduling parameter, a discrete-time linear parameter-varying model can be constructed, which is then used to develop a novel transition flight control architecture that integrates the adaptive model predictive control law with feedforward effect of the dynamic reference compensation. The simulation results demonstrate the effectiveness of proposed transition flight controllers and its robustness subject to external disturbance.

Sonic boom focus modeling of benchmark cases using the PCBoom suite

Focusing of sonic booms is a phenomenon that can arise from particular atmospheric conditions or aircraft trajectories including acceleration, turns, or other variations from steady level flight. Because such maneuvers are unavoidable, development of regulations for civil supersonic flight will depend on the capability to understand and predict the scenarios in which focus booms can lead to ground overpressures several times higher than typical carpet booms. Models for predicting sonic boom focal zones and focus boom signatures have been developed by several organizations, and a recent AIAA Sonic Boom Prediction Workshop provided an opportunity to compare results among such models. Rallabhandi et al. developed benchmark cases to further extend verification of focus boom models. The PCBoom suite of tools, including the lossy nonlinear Tricomi equation (LNTE) module, was used to predict caustic geometry and pressure signatures near the caustic arising from a constant acceleration level trajectory and also from an accelerating climb trajectory. A recently developed methodto streamline the focus boom modeling process within PCBoom was exercised as part of this benchmark case evaluation. [Work supported by the National Aeronautics and Space Administration under project VXAHA321.]

UH-60A Airloads Workshop—Setting the Stage for the Rotorcraft CFD/CSD Revolution, Part I: Background and Initial Success

The UH-60A Airloads Workshop was a unique collaboration of aeromechanics experts from the U.S. Government, industry, and academia to address technical issues that hindered accurate rotor loads predictions. The Airloads Workshop leveraged the NASA/Army UH-60A Airloads flight test and NFAC wind tunnel test data. It functioned continuously for 17 years, from 2001 to 2018, and brought about one of the most important advancements in rotorcraft aeromechanics prediction capabilities by successfully demonstrating high-fidelity coupled computational fluid dynamics (CFD) and computational structural dynamics (CSD) analyses for both steady and maneuvering flight. The article is divided into two parts. Part I surveys the background of rotorcraft CFD/CSD development difficulties, the origins of the Airloads Workshop, and the rapid success achieved during the first phase that consisted of eight Workshops. Part II describes ongoing development during the subsequent two phases of the Airloads Workshop, the Ninth through the 13th, and the 14th through the 31st Workshops; the impact of the Airloads Workshop; and the lessons learned. Part I surveys the technical activities that led to a breakthrough for CFD/CSD coupling to successfully predict rotor blade airloads in trimmed steady-level flight conditions. This success illustrated the importance of collaboration among key experts with diverse backgrounds focused on a common objective to advance rotorcraft prediction methods.

Development of a Flexible Framework Multi-Design Optimization Scheme for a Hand Launched Fuel Cell-Powered UAV

This paper presents different methods for the design of a hand-launchable, fixed wing, fuel cell-powered unmanned aerial vehicle (UAV) to maximize flight endurance during steady level flight missions. The proposed design methods include the development of physical models for different propulsion system components. The performance characteristics of the aircraft are modeled through empirical contributing analyses in which each analysis corresponds to an aircraft subsystem. The contributing analyses are collected to form a design structure matrix which is included into a multi-disciplinary analysis to solve for the design variables over a defined design space. The optimal solution is found using a comprehensive optimization tool developed for long endurance flight missions. Optimization results showed a significant improvement in UAV flight endurance that reached up to 475 min with take-off ratio equals to 59 min/kg. Wind tunnel and bench-top tests and HiL simulation tests are performed to validate the results obtained from the optimization tools. Validated optimization results showed an increase of the overall UAV flight endurance by 19.4% compared to classical approaches in design methods.

Minimum‐series twist distributions for yawing‐moment control during pure roll*

AbstractThe main wings of most aircraft produce adverse yaw during roll. In order to control the lateral direction of the aircraft during the roll, the rudder is often mixed with the aileron. Lifting‐line theory is used here to develop spanwise lift distributions that require the minimum number of terms in the Fourier‐series solution for controlling the yawing moment during pure rolling motion using only the main wing. It is shown that the yawing moment can be controlled for arbitrary rolling moments and/or rolling rates by adding symmetric twist in the main wing. The induced‐drag penalty for using this method to control the yawing moment is significant and discussed in detail. For example, it is shown that if zero yawing moment is prescribed, the induced drag can increase by 108% for a prescribed rolling moment or by 300% during a steady rolling rate relative to the induced drag in steady level flight. Because this is the minimum‐series solution, it does not represent the solution for yaw control with minimum induced drag, since more terms could be used in the Fourier series describing the lift distribution to control yaw with less induced drag. However, the solutions presented here can be useful for aircraft with continuous trailing‐edge technologies that are limited in spanwise deflection gradients, and provide a theoretical upper bound on the minimum induced‐drag that can be expected if the main wing is used for yawing‐moment control during pure roll.

Aerodynamic and stability analysis of a VTOL flying wing UAV

The stability analysis of an aerial vehicle is an of great importance integral part of its design procedure. It is of even greater importance in the case of tailless aircraft, which are prone to stability issues. In the present study, the aerodynamic and stability characteristics of a Vertical Take-off and Landing (VTOL) fixed wing Unmanned Aerial Vehicle (UAV), designated as MPU RX-4, are investigated. The MPU RX-4 has a flying wing layout and is capable of performing, both conventional flight, like a regular fixed wing aerial vehicle, as well as vertical hovering, like a multicopter, adapting on different operational demands and achieving rapid field deployment. In this study, the preliminary design phase of the MPU RX-4 is presented and the aerial vehicle’s aerodynamic performance, as well as, its stability and control behavior are assessed using both semi-empirical correlations, specifically modified for lightweight flying wing UAVs, and Computational Fluid Dynamics (CFD) analyses. These correlations are employed to estimate the non-dimensional aerodynamic coefficients for various flight conditions (e.g. cruise, loiter, maximum speed, etc.) of the MPU RX-4 flight envelope. Furthermore, the correlations are validated with dedicated CFD analyses in order to assure their level of accuracy. Finally, the MPU RX-4 stability and control derivatives, and the required control surfaces deflection for steady level flight are computed, in order to assess its overall aerodynamic performance and flight characteristics.

Influence of Propeller Location, Diameter, and Rotation Direction on Aerodynamic Efficiency

In this Paper, the integrated propeller–wing system design space is investigated to gain insight into the influence of propeller diameter, location, and rotation direction. To conduct this investigation, a methodology is developed that uses the propeller power required during steady-level flight as a metric for efficiency. Full mutual interaction between the propeller and wing are taken into account in the aerodynamic analysis. To ensure that the propeller is not operating off design, it is designed within an iterative trim loop. This approach is then used to investigate the theory that an inboard-up rotating propeller located at the wing tip is the most aerodynamically efficient configuration through the use of a test case. When considering the trimmed propeller power required, the trends for this case indicate that the optimal propeller for this planform is not an inboard-up rotating propeller at the wing tip, but instead is an inboard-down rotating propeller near the root of the wing.

Minimising induced drag with weight distribution, lift distribution, wingspan, and wing-structure weight

ABSTRACTBecause the wing-structure weight required to support the critical wing section bending moments is a function of wingspan, net weight, weight distribution, and lift distribution, there exists an optimum wingspan and wing-structure weight for any fixed net weight, weight distribution, and lift distribution, which minimises the induced drag in steady level flight. Analytic solutions for the optimum wingspan and wing-structure weight are presented for rectangular wings with four different sets of design constraints. These design constraints are fixed lift distribution and net weight combined with 1) fixed maximum stress and wing loading, 2) fixed maximum deflection and wing loading, 3) fixed maximum stress and stall speed, and 4) fixed maximum deflection and stall speed. For each of these analytic solutions, the optimum wing-structure weight is found to depend only on the net weight, independent of the arbitrary fixed lift distribution. Analytic solutions for optimum weight and lift distributions are also presented for the same four sets of design constraints. Depending on the design constraints, the optimum lift distribution can differ significantly from the elliptic lift distribution. Solutions for two example wing designs are presented, which demonstrate how the induced drag varies with lift distribution, wingspan, and wing-structure weight in the design space near the optimum solution. Although the analytic solutions presented here are restricted to rectangular wings, these solutions provide excellent test cases for verifying numerical algorithms used for more general multidisciplinary analysis and optimisation.

Performance Analysis of Medium Altitude Low-Cost Autonomous Quadcopter

In modern engineering, the rapid growth in the development of unmanned aerial vehicles which deserves the major space in the area of intelligence and surveillance operations. Unlike the conventional aircraft, the number of rotors mounted in an autonomous body that determines the propulsive efficiency as well as the maneuvering capabilities of unmanned aerial vehicles. The main objective of this research reveals that the low cost and an autonomous aerial vehicle with four rotors configuration design to achieve the long endurance flight limit. The developed Quadcopter flying model integrated with Arduino autopilot module which effectively controls the various maneuvering actions during the flight. This paper also reveals the entire hardware structure of the designed Quadcopter and its flight dynamic response characteristics. The remotely operated Quadcopter model was subjected to the flight test and the performance parameters have recorded. The flight test was conducted in an open environment in which stability parameters has been observed during onboard. Pitch angle, the Rotational speed of four rotors and the various attitudes of Quadcopter were changed accordingly to maintain the steady level flight. Findings are plotted with respect to the flight endurance.

Unsteady aerodynamic modeling and control of pusher and tilt-rotor quadplane configurations

Abstract A nonlinear unsteady aerodynamics model is coupled with a three degree of freedom quadplane to control the forward and backward transition between hover and steady level flight. The unsteady lift and drag forces are modeled using a lumped vortex model for flat plates. Two variants for the quadplane are considered: (i) a pusher and (ii) a tilt-rotor configuration in the absence of control surfaces to assess the controllability for altitude, attitude and forward speed. Conventional PID control is applied to generate the control inputs. The simulation results conclude that the pusher quadplane configuration is effortless to control as all the selected states are controllable, whereas for the tilt-rotor configuration, even though the vehicle is stable, altitude control is significantly more challenging due to one less control input when compared to the pusher configuration.

Analysis of low-speed height-velocity diagram of a variable-speed-rotor helicopter in one-engine-failure

This work analyzes the effects of variable operating rotor speed on the helicopter low-speed Height–Velocity (H-V) diagram in one engine inoperative (OEI) situation. Using the UH-60A helicopter as baseline helicopter, a flight dynamics model and the applied optimal control method were validated against the corresponding flight test data. Both low-speed H-V diagram in OEI and landing procedures at corresponding key points (Low hover, Knee and High hover) were investigated. Results indicate that the reduction of operating rotor speed will cause the unsafe area of H-V diagram to shrink gradually at first, and then expand rapidly when the operating rotor speed reduces below a certain value. It is observed that the lower constraint of rotor speed during OEI landing has a significant impact on the H-V diagram. Finally, results also show that a reasonable rotor speed not only effectively reduces the helicopter required power during steady level flight, but also improves the landing performance in the event of OEI.

Flight of an aeroplane with solid-state propulsion

Since the first aeroplane flight more than 100 years ago, aeroplanes have been propelled using moving surfaces such as propellers and turbines. Most have been powered by fossil-fuel combustion. Electroaerodynamics, in which electrical forces accelerate ions in a fluid1,2, has been proposed as an alternative method of propelling aeroplanes-without moving parts, nearly silently and without combustion emissions3-6. However, no aeroplane with suchasolid-state propulsion system has yet flown. Here we demonstrate that a solid-state propulsion system can sustain powered flight, by designing and flying an electroaerodynamically propelled heavier-than-air aeroplane. We flew a fixed-wing aeroplane with a five-metre wingspan ten times and showed that it achieved steady-level flight. All batteries and power systems, including a specifically developed ultralight high-voltage (40-kilovolt) power converter, were carried on-board. We show that conventionally accepted limitations in thrust-to-power ratio and thrust density4,6,7, which were previously thought to make electroaerodynamics unfeasible as a method of aeroplane propulsion, are surmountable. We provide a proof of concept for electroaerodynamic aeroplane propulsion, opening up possibilities for aircraft and aerodynamic devices that are quieter, mechanically simpler and do not emit combustion emissions.

Minimizing Induced Drag with Lift Distribution and Wingspan

Minimum induced drag for fixed gross weight and wingspan is obtained from the elliptic lift distribution. However, minimum induced drag for steady level flight is not obtained by imposing the constraints of fixed gross weight and wingspan. Because required wing-structure weight is a function of wingspan and lift distribution, there exist an optimum lift distribution and wingspan for a given weight and weight distribution that minimizes the induced drag in steady level flight. This optimum lift distribution can vary significantly from the elliptic lift distribution, depending on the design constraints. Analytic solutions for three such optimum lift distributions are presented for rectangular wings with varying sets of design constraints. These design constraints are 1) weight, maximum stress, and chord length fixed; 2) weight, maximum stress, and wing loading fixed; and 3) weight, maximum deflection, and wing loading fixed. It is shown that each of these optimum lift distributions results in lower induced drag than a fixed elliptic lift distribution for the same design constraints.

Development and Flight Testing of an Automatic Leveling Function for a Fighter Aircraft

In this paper a control topology with limited pitch and roll authority is described for automatic leveling of a fighter aircraft from disorientation. The proposed scheme provides recovery capability for an aircraft from any bank angle within the envelope defined by airspeed (CAS), angle of attack (AoA) and flight path angle to steady level flight. The control algorithm generates pitch and roll commands to bring the aircraft to level flight without violating the specified aircraft parameter limits. The command generation scheme is validated through nonlinear simulation and flight tested on a fighter aircraft for different flight conditions and the results are presented.