Aircraft Velocity Research Articles

Flight test for the noise source modeling of the optionally piloted personal air vehicle

An acoustic flight test is carried out in order to obtain the noise hemisphere of the optionally piloted personal air vehicle (OPPAV). The OPPAV is one-seater, research urban air mobility (UAM) being developed by Korea Aerospace Research Institute. Noise radiated from the aircraft is measured using a microphone array consisting of a total of 78 microphones. Flight logs such as the aircraft position, velocity, and attitude are recorded synchronously with the acoustic signals. Measured acoustic signals are dedopplerized and divided into a number of segments. Noise source model for the OPPAV flight is constructed using these segments. The noise source model will be used for the assessment of noise annoyance during urban flight of the UAMs.

Off-design analysis of the inverted Brayton cycle engine

PurposeThe purpose of this study is to perform an off-design analysis of the inverted Brayton cycle engine.Design/methodology/approachThe off-design analysis equations of the inverted Brayton cycle engine were first derived in this study and the control parameters of the inverted Brayton cycle engine were first determined and investigated.FindingsIt is observed that by controlling the total temperature decrease in cooling section, it is possible to adapt the engine for low specific fuel consumption working conditions or high thrust working conditions. Specific fuel consumption is reduced by 27.1 % by stopping cooling in the cooling section and thrust is increased by 27.6 % by working with full load of the cooling section (500 K temperature decrease in cooling section). It is observed that thrust depending on the flight Mach number increases with an increase in flight Mach number and reaches a peak value at 5.21 flight Mach number and reduces by 80.8 % at 6 flight Mach number relative to the peak value. The specific fuel consumption increases rapidly as the Mach number increases, and the specific fuel consumption is 49.0 g/[kN.s] at Mach 1, reaches 70.4 g/[kN.s] at Mach 5 and increases to 412 g/[kN.s] at Mach 6. The specific fuel consumption increases from 68.1 to 73.0 g/(kN.s) and the thrust decreases from 16.5 to 13.3 kN as the total preburner exit temperature increases from 1,500 to 2,000 K. Specific fuel consumption decreases from 83.1 to 64.8 g/(kN.s) and thrust increases from 4.60 to 11.08 kN depending on afterburner exit total temperature increase from 1,800 to 2,500 K.Research limitations/implicationsThe cooling section reduces total temperature of the gas flow to lower values to increase the compressor total pressure ratio. The compressor increases the total pressure of the gas flow to the optimum total pressure ratios to increase the nozzle exit Mach number and gain more thrust. The afterburner increases the total temperature of the gas flow to increase the sound speed in the nozzle exit to increase thrust. The nozzle expands the gas flow to reduce the static pressure of the gas flow to near the optimum value, atmosphere pressure, to increase thrust and reduce specific fuel consumption.Practical implicationsHypersonic and supersonic air vehicles can use the current engine model for the its own propulsion systems.Social implicationsAfter first heavier than air flight, aero engines was designed for only used for aero vehicle. Internal combustion engines were used for propelled propeller aircraft at the first term of aircraft. However, propeller-propelled aircrafts are not sufficient to increase aircraft velocity to supersonic Mach numbers due to the shock losses of propeller, so the supersonic era was only introduced by revolution in propulsion systems with new concept. A jet engine was developed to be candidate for supersonic flight.Originality/valueOff-design analysis equations of an inverted Brayton cycle engine were first derived in this study. Furthermore, the control parameters of the inverted Brayton cycle engine were first determined and investigated in this paper.

ON-BOARD CONTROL-MEASUREMENT SYSTEM FOR MICRO CONVERTIPLANE-TYPE UNMANNED AERIAL VEHICLES

The article describes the on-board control-measurement system designed to record the main flight parameters of a micro-convertiplane type unmanned aerial vehicle and the element base used for this purpose. The methodology of conducting tests in stationary mode through the on-board control-measurement system was developed and tests were carried out. The engine's diagnostic check and condition assessment has been performed using the on-board control-measuring system. The possibility of on-ground check of the micro unmanned aerial vehicles power elements before and after the flight, as well as the possibility of monitoring the flight events that may occur in the air during the flight has been shown. The data on the power consumption and temperature of each engine in the stationary mode, the direction of the aircraft axes and the velocity vector have been recorded by mean of the developed on-board control-measurement system, as well as time dependence diagrams of the corresponding parameters have been built in comparative order based on the obtained values. A methodology has been developed for the diagnostics of the engine operation of the micro-unmanned aerial vehicle before and during the flight. It has been shown that the temperature, power consumption and number of revolutions of the engines included in the power system of the aircraft, as well as the aircraft axes and velocity vector direction data is recorded during the flight, by mean of the on-board control-measurement system designed to collect data on the flight-technical parameters of the convertiplane-type micro-unmanned aerial vehicles.

Characterization of the airborne aerosol inlet and transport system used during the A-LIFE aircraft field experiment

Abstract. Atmospheric aerosol particles have a profound impact on Earth's climate by scattering and absorbing solar and terrestrial radiation and by impacting the properties of clouds. Research aircraft such as the Deutsches Zentrum für Luft- und Raumfahrt e.V. (DLR) Falcon are widely used to study aerosol particles in the troposphere and lower stratosphere. However, transporting a representative sample to the instrumentation inside the aircraft remains a challenge due to high airspeeds and changing ambient conditions. In particular, for high-quality coarse-mode aerosol measurements, knowledge about losses or enhancements in the aerosol sampling system is crucial. In this study, the sampling efficiency of the aerosol inlet aboard the Falcon research aircraft is characterized for the first time with state-of-the art in situ measurements including sizing instruments operated behind the Falcon aerosol inlet and mounted at the aircraft wing not affected by the aerosol inlet. Sampling efficiencies were derived for different true airspeed ranges by comparing the in-cabin and ”full”-size-range particle number size distributions during 174 flight sequences with a major contribution of mineral dust particles during the ”Absorbing aerosol layers in a changing climate: aging, lifetime and dynamics” project (A-LIFE). Additionally, experimentally derived Stokes numbers were used to calculate the cutoff diameter of the aerosol sampling system for different particle densities as a function of true airspeed. As expected, the results show that the velocity of the research aircraft has a major impact on the sampling of coarse-mode aerosol particles with in-cabin instruments. For true airspeeds up to about 190 m s−1, aerosol particles larger than about 1 µm are depleted in the sampling system of the Falcon during the A-LIFE project. In contrast, for true airspeeds higher than 190 m s−1, an enhancement of particles up to a diameter of 4 µm is observed. For even larger particles, the enhancement effect at the inlet is still present, but inertial and gravitational particle losses in the transport system get more and more pronounced, which leads to a decreasing overall sampling efficiency. In summary, aerosol particles are either depleted or enhanced in the Falcon aerosol inlet, whereas transport in sampling lines always leads to a loss of particles. Here, we have considered both effects and determined the cutoff diameter for the A-LIFE transport system (i.e., the sampling lines only), the cutoff diameter of the Falcon aerosol inlet (i.e., the effect of the inlet only), and the combined effect of the inlet and sampling lines.

Swarm Division-Based Aircraft Velocity Obstacle Optimization Considering Low-Carbon Emissions

In the pursuit of sustainable aviation, this paper presents an innovative approach that adopts a swarm division strategy to enhance and refine the velocity obstacle (VO) method, guided by a low-carbon principle. A dynamic elliptical protection zone model forms the core of this innovative approach. Specifically, this dynamic elliptical protection zone is created based on the difference in aircraft velocity, and a swarm division strategy is introduced in this process. Initially, aircraft that share the same route and type, and have similar velocities and distances, are grouped into swarms. Then, the characteristics of the swarms, such as mass points, velocities, and protection zones, are recorded. Second, the collision cone (CC) between swarms is established, and planar geometrical analysis is used to determine the optimal relief velocity and heading of aircraft on the low-carbon objective while ensuring a safe interval between aircraft in the swarm during the relief period. Additionally, a swarm control algorithm is utilized to adjust the velocity of the aircraft by a small margin. Finally, simulation experiments are conducted using Python, revealing that the swarm relief efficiency of the enhanced VO method sees a notable increase of over 33%. Concurrently, the need for adjustments decreases by an average of 32.78%, while fuel savings reach as high as 70.18%. The strategy is real-time and operational, significantly reduces the air traffic controller (ATC) workload, improves flight efficiency and safety, and contributes positively to the reduction in carbon emissions, which is beneficial for the environment.

Flutter Phenomenon and Safety Implications in Transonic Flow

As advancements in aircraft manufacturing technology persistently elevate the boundaries of flight speed, challenges emerge when aircraft velocities verge on sonic speeds. Notably, empirical observations have highlighted instances of pronounced vibrations and, in extreme cases, structural failures, casting shadows over flight stability and safety. This research endeavors to dissect the ramifications of shock waves and flutter phenomena, precipitated by fluid discontinuities during supersonic flight, on the overarching safety of aircraft. Methodologically, the study delineates a series of simulation protocols, commencing with the derivation of simplified functions representing the aircraft's 2D contour, followed by error function computations, and finally, using the finite difference method in conjunction with the successive over-relaxation (SOR) iterative technique to model the aircraft's velocity distribution. Preliminary findings indicate that for β values spanning 0.1, 0.15, and 0.175, the surrounding airflow retains its continuity and stability, with the Mach number's rate of change exhibiting a decelerating trend. Conversely, at a β threshold of 0.2, the airflow descends into turbulence, manifesting in erratic Mach number fluctuations. Such fluidic discontinuities underscore potential threats to the aircraft's flight safety, necessitating further exploration and mitigation strategies.

Design charts for hydraulic twister based aircraft arrester gear

In this article, arrestment of an aircraft by a hydraulic twister based system is modeled and reduced to simple design charts. We are primarily interested in performance parameters like maximum deceleration, run-out distance, and maximum tape force. We first present a simplified mechanical model for the system. Some analytical progress is possible with the governing nonlinear differential equation, yielding expressions for velocity and acceleration in terms of displacement. The system has only two nondimensional parameters. This allows us to present simple design charts for the performance parameters. The same approach is extended to secondary performance parameters namely maximum jerk and the time at which maximum deceleration occurs. We choose some practical parameter sets and demonstrate the use of our charts. These charts can be used for design calculations and quick performance prediction of an arresting gear design for different weight classes of aircraft and practical initial velocities.

Impact Analysis of a Commercial Airplane Over a Nuclear Spent Fuel Dry Storage Cask

The evaluation of the structural integrity of a vertical cask that is used for spent nuclear fuel dry storage is reported. The cask diameter and height are 3.566 m (140 in.) and 5.28 m (207.75 in.), respectively. The analysis focuses on such a cask being impacted by a commercial airplane. The dry storage container standards, which are under evaluation and approval by the U.S. Nuclear Regulatory Commission, are considered. The storage container inner basket is made of a stainless steel plate cylinder. It is located within an outer shell. The last one is manufactured with concrete and has internal and external steel liners. The commercial airplane considered in this analysis has a length of 40.39 m (132 ft, 6 in.). Its wingspan and height are 35.23 m (115 ft, 7 in.) and 11.98 m (39 ft, 4 in.), respectively. Its take-off weight is 81 090 kg (178 773 lb). An explicit analysis with the finite element method is carried out. The impact angles were 0, 30, 45, and 60 deg with respect to the horizontal. The mesh of the domain has 1 104 229 hexahedral elements and 1 516 156 nodes. Initially, all the structures are considered without restrictions and free of stresses. The vertical container for dry storage is at rest on a rigid concrete base. The aircraft velocity is 234 m/s or 842 km/h (523 mph). The impact event is simulated in an interval of 0.03 s. The maximum principal stress fields show that there are points at the lid of the container that are above the elastic limit and the ultimate strength. Under these conditions, brittle failure is expected.

Microcontroller Based Aeronautical Crash Prevention System through RF Communication

Embedded system is a special-purpose computer, designed to perform a dedicated function. Embedded system is quickly growing in various fields like industrial automation, home appliances, automobiles, aeronautics etc. The main aim of this project is to prevent the collision among air flights. Many methods are available to prevent the collision between the flights, but here RF technology and Microcontroller is used which provides some additional features for future enhancement. The system uses a microcontroller to collect data from sensors on the aircraft and process it to determine the distance and velocity of other aircraft in its vicinity. The system then sends signals to other aircraft in its range through RF communication to alert them about the potential danger. In addition performance evaluation of the system, which includes the detection accuracy, response time, and communication reliability, is monitored.

The High-speed Aircraft Trajectory Prediction Method Based on Deep Learning Network

This paper applies deep learning technology to the trajectory prediction research of high-speed aircraft in order to predict the medium- and long-term trajectory. Hence, a hybrid model combining the time series convolution neural network (TSCNN) and the long short-term memory (LSTM) is put forward to predict position and velocity of high-speed aircraft. In the TSCNN, convolution operations are used to extract trajectory features of time and space to predict the future trajectory. However, the error will be generated during prediction the process, so the LSTM network can be tapped to correct it by analyzing why it is produced in a bid to address error accumulation. According to simulations conducted, the model can lead to high-precision prediction.

Research on airspace sector optimization based on Voronoi diagram and improved K-means algorithm

Sector partition is an important task of air traffic control, and a reasonable sector partition can improve the utilization rate of airspace and protect the flight safety of aircrafts. Since the sector partition during flat hours is not well suited to the complex air situation, this paper proposes a sector optimization method based on Voronoi diagram and improved K-means. Firstly, a conflict network is constructed based on the air situation, and a comprehensive sector control workload measurement method is proposed by combining aircraft velocity obstacle relationship and complex network theory. Based on the workload value, a cluster center is determined as the generating element of Voronoi diagram by using the improved K-means method, and then the sector is optimized by using the division method of Voronoi diagram. In this paper, the data of Xiamen airspace control sectors are collected as a simulation scenario for calculation and analysis. The simulation results show that the average variance of the optimized sector control workload is reduced by 66.04% during the peak hours and 13.88% during the flat hours compared with the original sector. The method achieves the purpose of balancing the sector workload, verifies the effectiveness of the sector optimization method, and provides a reference basis for the existing sector partition work.

Dilution of Precision in the Problem of Estimating the Aircraft Velocity by Means of Three-Beam Lidar Measurements

Dilution of Precision in the Problem of Estimating the Aircraft Velocity by Means of Three-Beam Lidar Measurements

Modelling of spinal injury during aircraft seat ejection loads

The primary aim of the present work is to calculate the ejection forces on the spine lumbar vertebra (L1-L5) using Finite element modeling. The critical time of ejection was calculated based on mathematical modelling using ejection parameters such as tail height, aircraft velocity, and aircraft length. The Finite element analysis at a different level of G-Forces was performed with appropriate boundary conditions to simulate seat ejection and stress on the vertebral body is analyzed. The model was validated at the Functional spinal unit level and a multi-level model based on published literature. It was found that the anterior side of the L1 Vertebrae experiences the maximum stress, which justifies the common anterior wedge compression fracture. More stress was distributed on the superior region of endplates than that at the inferior region throughout the lumbar vertebra although the difference in the stress value between the superior and inferior region decreases from L1 to L5. In all the cases cortical bone experiences the maximum stress followed by the endplate and the stress value obtained for Intervertebral Disc(IVD) is found to be 90% lesser than the endplate, which means the initial failure may occur at the vertebral body followed by endplate.

Dynamic Aircraft Wake Separation Based on Velocity Change

Traditional research on static wake-vortex reduction usually considers only the influence of external environmental factors, while ignoring the dynamic change in an aircraft’s flight state. In order to solve this problem, this paper proposes a method to reduce separation using dynamic wake information based on changes in flight velocity. Firstly, relying on the wake-vortex generation and dissipation model, the initial circulations and dissipation parameters of the wake vortex at different aircraft velocities are calculated. Then, the complete evolution process of the wake vortex generated by different types of aircraft at different velocities is analyzed, and the evolution law of a wake vortex with changing velocity is obtained. Afterwards, according to the actual aerodynamic shape of CRJ-900, the aerodynamic model for CRJ-900 when it encounters a wake vortex is established. The situation of an CRJ-900 encountering a wake vortex under TBS is analyzed, which proves that TBS is safe and has a certain reduction potential. Then, taking the rolling moment coefficient as the safety index, the minimum safe separations at different velocities are calculated. Finally, a simulation for the separation reduction based on a dynamic wake vortex is carried out. Compared with the static wake-based separation, the dynamic wake-based separation technology can greatly reduce the aircraft separation requirements while ensuring the operation safety. The final simulation results show that the dynamic separation reduction in CRJ-900 following a medium and heavy aircraft can reach 44.3% and 51.6%, respectively.

Hierarchical Reinforcement Learning Approach for Autonomous Cross-Country Soaring

Solving the decision-making problem between pursuing the objective of covering distance and exploiting thermal updrafts is the central challenge in cross-country soaring flight. The need for trading short-term rewarding actions against actions that pay off in the long term makes for a hard-to-solve problem. Policies resulting from reinforcement learning offer the potential to handle long-term correlations between actions taken and rewards received. The paper presents a reinforcement learning setup, which results in a control strategy for the autonomous soaring sample application of GPS Triangle racing. First, we frame the problem in terms of a Markov decision process. In particular, we present a straightforward model for the three-degrees-of-freedom system dynamics of a glider aircraft that does not make any simplifying assumptions regarding the wind field or the relative aircraft velocity. The competition task is decomposed into subtasks, then. Stochastic gradient ascent solves the associated hierarchical reinforcement learning problem without the designer employing any further, potentially deficient heuristics. We present an implementation of the overall policy alongside an updraft estimator on embedded hardware aboard an unpiloted glider aircraft. Flight-test results validate the successful transfer of the hierarchical control policy trained in simulation to real-world autonomous cross-country soaring.

Aircraft trajectory generation and control for minimum fuel and time efficient climb

This paper presents a novel approach for generating the best possible climb trajectory that ensures minimum fuel and time efficient climb. The problem is first formulated using standard steady climb equations, which generate a unique combination of flight velocity and flight path angle at each altitude. A possible scenario, such as air density, mass, available power, and required powered variation with altitude, is taken into account when defining the problem. Thereafter, sliding-mode-based trajectory tracking control is formulated with its design procedures, system stability with applied control inputs, finite-time convergence analysis, and complete architecture. A Hansa-3 research aircraft is considered as an example model to demonstrate the work. The findings of generated trajectory are then produced and discussed. In order to follow the design trajectory and achieve the same, the sliding-mode-based control command is supplied. The novelty of the present work lies in proposed strategy of trajectory generation, wherein the aircraft path and velocity are found out to make the fuel and time efficient climb possible. Subsequently, robust control law is developed which shows the applicability of the proposed work on autopilot. The results show that the proposed controller not only controls the aircraft but is also able to follow the design trajectory with minimal errors. To further explore the impact of aircraft mass on climb performance, repeated set simulation is carried out. The outcome is compared with conventional climb, which promises its practical implementation since the proposed solution is simple and compatible to integrate with the existing aircraft autopilot.

Robust aircraft conflict resolution under trajectory prediction uncertainty

In air traffic control, aircraft velocity may be perturbed due to weather effects or measurements errors and affect trajectory prediction. We address the aircraft conflict resolution problem in the presence of such perturbations. We consider polyhedral uncertainty sets on aircraft velocity, develop a budgeted robust optimization approach and show that it is amenable to mixed-integer optimization. Numerical experiments reveal that perturbations of the order of 5% on aircraft velocities can be accounted for without significantly impacting performance.

Analytical Aircraft State and IMU Signal Generator From Smoothed Reference Trajectory

Generating aircraft position, velocity, and attitude along a reference trajectory is useful in many path and mission planning applications. In applications simulating an inertial navigation system the body-frame accelerations and angular rates are also required. This work presents a method for generating a smoothed continuous-curvature reference trajectory from a series of waypoints using line, arc, clothoid, and Fermat spiral segments in a fillet smoothing framework. The geometry of the segments are used to generate the position, velocity, and attitude of an aircraft following the reference trajectory. A primary contribution of this article is the use of the path geometry, coordinated turn model, and curvilinear motion theory to obtain analytical solutions for the body-frame accelerations and angular rates of the aircraft along the reference trajectory.

SDC-based dynamic inversion controller for a fixed-wing aircraft

This work presents dynamic model inversion controller which uses state dependent coefficient matrices, and is applied for a fixed-wing aircraft attitude and velocity control problem. Numerical validation of the proposed controller is done using a Cessna 172 aircraft simulation model for several scenarios of atmospheric disturbances.

An approximate dynamic programming approach for solving an air combat maneuvering problem

Within visual range air combat involves execution of highly complex and dynamic activities, requiring rapid, sequential decision-making to achieve success. Fighter pilots spend years perfecting tactics and maneuvers for these types of combat engagements, yet the ongoing emergence of unmanned, autonomous vehicle technologies elicits a natural question — can an autonomous unmanned combat aerial vehicle (AUCAV) be imbued with the necessary artificial intelligence to perform challenging air combat maneuvering tasks independently? We formulate and solve the air combat maneuvering problem (ACMP) to examine this important question, developing a Markov decision process (MDP) model to control a defending AUCAV seeking to destroy an attacking adversarial vehicle. The MDP model includes a 5-degree-of-freedom, point-mass aircraft state transition model to accurately represent both kinematics and energy while maneuvering. An approximate dynamic programming (ADP) approach is proposed wherein we develop and test an approximate policy iteration algorithm that implements value function approximation via neural network regression to attain high-quality maneuver policies for the AUCAV. A representative intercept scenario is specified for testing purposes wherein the AUCAV must engage and destroy an adversary aircraft attempting to penetrate the defended airspace. Several designed experiments are conducted to determine how aircraft velocity and adversary maneuvering tactics impact the efficacy of the proposed ADP solution approach and to enable efficient algorithm parameter tuning. ADP-generated policies are compared to two benchmark maneuver policies constructed from two reward shaping functions found in the ACMP literature, attaining improved mean probabilities of kill for 24 of 36 air combat situations considered.