Cruise Aircraft Research Articles

Large Eddy Simulation for Empirical Modeling of the Wake of Three Urban Air Mobility Vehicles

Recent advances in urban air mobility have driven the development of many new vertical take-off and landing (VTOL) concepts. These vehicles often feature original designs departing from the conventional helicopter configuration. Due to their novelty, the characteristics of the supervortices forming in the wake of such aircraft are unknown. However, these vortices may endanger any other vehicle evolving in their close proximity, owing to potentially large induced velocities. Therefore, improved knowledge about the wakes of VTOL vehicles is needed to guarantee safe urban air mobility operations. In this work, we study the wake of three VTOL aircraft in cruise by means of large eddy simulation. We present a two-stage numerical procedure that enables the simulation of long wake ages at a limited computational cost. Our simulations reveal that the wakes of rotary vehicles (quadcopter and side-by-side helicopter) feature larger wake vortex cores than an isolated wing. Their decay is also accelerated due to self-induced turbulence generated during the wake roll-up. A tilt-wing wake, on the other hand, is moderately turbulent and has smaller vortex cores than the wing. Finally, we introduce an empirical model of the vortex circulation distribution that enables fast prediction of wake-induced velocities, within a 2% error of the simulation results on average.

Estimating the noise dose range of the NASA X-59 aircraft in supersonic cruise using PCBoom propagation simulations.

PCBoom propagation simulations of the NASA X-59 low-boom in supersonic cruise were conducted using near-field CFD pressure waveforms as inputs to estimate the noise dose range. Low-booms were propagated through realistic atmospheric profiles from the Climate Forecast System Version 2 across the USA, and loudness statistics of the low-booms are presented. The near-field waveforms correspond to aircraft configurations expected to produce minimum- and maximum-loudness levels on the ground. The simulations show that the practical Perceived Level noise dose range is 72 to 86.7 dB. Dose ranges for other single event and cumulative sonic boom metrics of interest are also provided.

Time-Fuel-Optimal Navigation of a Commercial Aircraft in Cruise With Heading and Throttle Controls Using Pontryagin’s Maximum Principle

In this research, we consider the commercial aircraft trajectory optimization problem for a general cruise model with arbitrary spatial wind fields to be solved using the Pontryagin’s maximum principle. The model features two fundamental controls, namely “throttle setting” (appearing as a singular control) and “heading angle” (appearing as a regular control). For a problem with state-inequality constraints and minimum time-fuel objective, we show that the optimal “heading angle” can be described through the classic Zermelo’s navigation identity. We also demonstrate, by analyzing the switching function, that the singular “throttle setting” can be characterized through a feedback function that relies on both the optimal states and “heading angle”. The switching-point algorithm is employed to solve a case study where we inspect the optimality conditions and graph the optimal controls together with the optimal state and co-state variables.

Development of an Adaptive Aero-Propulsive Performance Model in Cruise Flight – Application to the Cessna Citation X

To accurately predict the amount of fuel needed by an aircraft for a given flight, a performance model must account for engine and airframe degradation. This paper presents a methodology to identify an aero-propulsive model to predict the fuel flow of an aircraft in cruise, while considering initial modeling uncertainties and performance variation over time due to degradation. Starting from performance data obtained from a Research Aircraft Flight Simulator, an initial aero-propulsive model was identified using different estimation methods. The estimation methods studied in this paper were polynomial interpolation, thin-plate splines, and neural networks. The aero-propulsive model was then structured using two lookup tables: one lookup table reflecting the aerodynamic performance, and another table reflecting the propulsive performance. Subsequently, an adaptative technique was developed to locally and then globally, adapt the lookup tables defining the aero-propulsive model using flight data. The methodology was applied to the Cessna Citation X business jet aircraft, for which a highly qualified level D research aircraft flight simulator was available. The results demonstrated that by using the proposed aero-propulsive performance model, it was possible to predict the aerodynamic performance with an average relative error of 0.99%, and the propulsive performance with an average relative error of 3.38%. These results were obtained using the neural network estimation method.

Conceptual design and its optimization of an air-cooled nuclear reactor for aircraft propulsion considering both thermal and neutronic performance

Air-cooled nuclear-powered engine is a good choice for low-attitude and high-speed aircraft soon. To achieve better propulsion performance, a conceptual design for air-cooled reactors with the aim of minimum size and high thrust-weight ratio is conducted in this work, where both thermal-hydraulics and neutronics targets are satisfied to make sure its safety under extreme operating conditions. The nuclear engine composed of the diffuser, the reactor, and the nozzle is modeled with simplification to mainly calculate heat transfer and pressure drop inside the flow channels of the reactor core. The relation of thrust and specific impulse is investigated to show the operation range of the engine. Geometrical dimensions of the air-cooled reactor core are minimized to offer the rated output thrust of 120 kN. For this optimal design, wall temperature of the reactor remains below the upper limit of 1644 K, and the core theoretically reaches criticality based on the critical buckling got from Monte-Carlo simulations of a fuel unit. Additionally, the neutronics analysis is then employed for the entire reactor. Attribute to Gd2O3 added in the fuel matrix, shutdown margin keeps below −1 $ under the submersion accident. This work finally proposes a design of the air-cooled reactor with a mass of 2.6 ton loaded by high-enriched uranium oxide, which could provide the high-speed cruise aircraft with the thrust of 120 kN at sea level for 15 days.

Enhancing the take-off performance of hypersonic vehicles using the improved chimp optimisation algorithm

AbstractThe performance of hypersonic vehicles in the take-off stage considerably influences their capability of accomplishing the flight tasks. This study is aimed at enhancing the take-off performance of a cruise aircraft using the improved chimp optimisation algorithm. The proposed algorithm, which uses the Sobol sequence for initial population generation and a function of the weight factors, can effectively overcome the problems of premature convergence and low accuracy of the original algorithm. In particular, the Sobol sequence aims to obtain a better fitness value in the first iteration, and the weight factor aims to accelerate the convergence speed and avoid the local optimal solution. The take-off mass model of the hypersonic vehicle is constructed considering the flight data obtained using the pseudo-spectral method in the climb phase. Simulations are performed to evaluate the algorithm performance, and the results show that the algorithm can rapidly and stably optimise the benchmark function. Compared to the original algorithm, the proposed algorithm requires 28.89% less optimisation time and yields an optimised take-off mass that is 1.72kg smaller.

Aircraft Mathematical Model Identification for Flight Trajectories and Performance Analysis in Cruise

In this paper, a mathematical model for estimating the performance and flight trajectories in cruise is identified from data available in flight manuals. The first part of this paper focuses on the design of a fuel flow and emissions model. Starting from the equations of motion of an aircraft in cruise, a simplified model representing the fuel flow in a corrected form was developed. A practical algorithm was next developed to identify the aircraft model parameters and to determine the mathematical structure that reflects its fuel flow. This process was done using performance data available in the aircraft flight crew operating manual. The emissions model was also developed based on data available in the International Civil Aviation Organization’s engine emissions databank. The second part of the paper deals with the development of algorithms for predicting the trajectories and calculating the optimal speeds (i.e., maximum range, long range, and economy) of the aircraft in cruise. Practical techniques for storing and retrieving information without using optimization algorithms have been considered. The methodology was applied on both a Cessna Citation X business jet and Bombardier CRJ-700 regional jet aircraft. The comparison results showed a very good agreement for the fuel consumption and optimal speed.

Study on Trajectory Optimization of Hypersonic Vehicle Based on Neural Network

For the horizontal take-off hypersonic cruise aircraft, research on the combined design method of multi-section was carried out, the main design parameters of different sections were analyzed, the parametric design model of the flight path was established, and the characteristics of the typical flight path were studied. On this basis, the calculation of sample points was carried out, and a prediction model of aircraft range and flight time based on the design parameters of the four main flight sections was established based on the neural network method. The genetic algorithm is used to optimize the flight path of the prediction model with the range as the objective function. The research results show that the neural network prediction model based on the parametric design of the trajectory can predict random sample points better than the trajectory model For the prediction of random sample points, compared with the calculation results of the trajectory model, the maximum errors of the flight range and flight time are within 0.82% and 0.45%. The prediction model is optimized with the flight range as the objective function, and the relative error between the optimal range and the trajectory model under the corresponding section parameters is less than 0.2%, which shows that the model established in this paper can better predict the range and flight time according to the section design parameters. Parametric modeling and neural network optimization are feasible methods for aircraft trajectory design and section parameter optimization.

Spiral instability modes on rotating cones in high-Reynolds number axial flow

This work shows the behavior of an unstable boundary-layer on rotating cones in high-speed flow conditions: high Reynolds number Rel>106, low rotational speed ratio S<1–1.5, and inflow Mach number M = 0.5. These conditions are most-commonly encountered on rotating aeroengine nose cones of transonic cruise aircraft. Although it has been addressed in several past studies, the boundary-layer instability on rotating cones remains to be explored in high-speed inflow regimes. This work uses infrared-thermography with a proper orthogonal decomposition approach to detect instability-induced flow structures by measuring their thermal footprints on rotating cones in high-speed inflow. The observed surface temperature patterns show that the boundary-layer instability induces spiral modes on rotating cones, which closely resemble the thermal footprints of the spiral vortices observed in past studies at low-speed flow conditions: Rel<105, S > 1, and M≈0. Three cones with half-cone angles ψ=15°, 30°, and 40° are tested. For a given cone, the Reynolds number relating to the maximum amplification of the spiral vortices is found to follow an exponential relation with the rotational speed ratio S, extending from the low- to high-speed regime. At a given rotational speed ratio S, the spiral vortex angle appears to be as expected from the low-speed studies, irrespective of the half-cone angle.

Thrust Vectoring of a Fixed Axisymmetric Supersonic Nozzle Using the Shock-Vector Control Method

The application of the Shock Vector Control (SVC) approach to an axysimmetric supersonic nozzle is studied numerically. SVC is a Fluidic Thrust Vectoring (FTV) strategy that is applied to fixed nozzles in order to realize jet-vectoring effects normally obtained by deflecting movable nozzles. In the SVC method, a secondary air flow injection close to the nozzle exit generates an asymmetry in the wall pressure distribution and side-loads on the nozzle, which are also lateral components of the thrust vector. SVC forcing of the axisymmetric nozzle generates fully three-dimensional flows with very complex structures that interact with the external flow. In the present work, the experimental data on a nozzle designed and tested for a supersonic cruise aircraft are used for validating the numerical tool at different flight Mach numbers and nozzle pressure ratios. Then, an optimal position for the slot is sought and the fully 3D flow at flight Mach number M∞=0.9 is investigated numerically for different values of the SVC forcing.

Plasma actuation effect on a NACA 4412 airfoil

Purpose This paper aims to investigate the effects of a dielectric barrier discharge (DBD) plasma actuator (PA) qualitatively on aerodynamic characteristics of a 3 D-printed NACA 4412 airfoil model. Design/methodology/approach Airflow visualization study was performed at a Reynolds number of 35,000 in a small-scale open-loop wind tunnel. The effect of plasma actuation on flow separation was compared for the DBD PA with four different electrode configurations at 10°, 20° and 30° angles of attack. Findings Plasma activation may delay the onset of flow separation up to 6° and decreases the boundary layer thickness. The effects of plasma diminish as the angle of attack increases. Streamwise electrode configuration, in which electric wind is produced in a direction perpendicular to the freestream, is more effective in the reattachment of the airflow compared to the spanwise electrode configuration, in which the electric wind and the free stream are in the same direction. Practical implications The Reynolds number is much smaller than that in cruise aircraft conditions; however, the results are promising for low-velocity subsonic airflows such as improving control capabilities of unmanned aerial vehicles. Originality/value Superior efficacy of spanwise-generated electric wind over streamwise-generated one is demonstrated at a very low Reynolds number. The results in the plasma aerodynamics literature can be reproduced using ultra-low-cost off-the-shelf components. This is important because high voltage power amplifiers that are frequently encountered in the literature may be prohibitively expensive especially for resource-limited university aerodynamics laboratories.

Flight management system for hydrogen-powered aircraft in cruise

The minimization of the Direct Operating Cost (DOC) for hydrogen-powered aircraft is formulated in this paper as an optimal control problem and is solved based on Pontryagin's minimum principle. As a consequence, the optimum cruise flight speed is determined assuming cruising at a constant altitude. The optimization criterion corresponds to the minimization of a functional representing the trade-off between the cost of hydrogen fuel and time-dependent costs, which are related by a parameter denoted by cost index. The value of this parameter is introduced by a pilot into the Flight Management System (FMS) of the aircraft. The HY4 aircraft model is used to obtain numerical results for the proposed methodology.

Impact of lift propeller drag on the performance of eVTOL lift+cruise aircraft

Electric vertical takeoff and landing (eVTOL) aircraft of the lift+cruise configuration are significantly impacted by the drag of the lift propellers even though they are only briefly used for takeoff and landing. This work aims to quantify the drag penalty of these high lift propellers, its impact on the performance of a lift+cruise eVTOL aircraft, and potential benefits from retracting the propellers during cruise. An eVTOL model is tested in the wind tunnel with propellers extended and retracted. Results show that retracting the propellers reduces parasitic drag by 38%. This measured drag reduction is used to estimate the performance of a passenger eVTOL and a surveillance drone employing a propeller retraction system. A passenger carrying eVTOL lift+cruise aircraft with a propeller retraction system could cruise 21% faster, at the same mission range, or increase range by 13% at the same flight speed. When endurance is the main mission objective, the benefits are smaller, in the order of 7%, and the potential gains are more sensitive to the weight of the retraction system.

Analysis and comparison of potential power and thermal management systems for high-speed aircraft with an optimization method

• An advanced PTMS with an indirect regenerative air cycle is analyzed. • A comparison based on minimum fuel weight penalty optimization is proposed. • Optimal matching of the PTMS and flight mission is achieved at Mach 1-4.4. • The PTMS with a ram air bleed mode is suitable for Mach 3.4–4.4 cruise aircraft. Under the dual effects of aerodynamic heating and high-power electronic equipment heating, the heat sink and power demand of advanced high-speed aircraft have been exponentially rising, which seriously restricts the aircraft performance. To improve system cooling and power supply performance and reduce engine performance loss, a power and thermal management system (PTMS) with high performance, low energy consumption, and light weight urgently needs to be developed. In this paper, three modes of a potential PTMS with different heat sinks and bleed air sources are further discussed to analyze and compare the optimal matching with the flight mission at Mach 1–4.4. The equivalent mass method is used to uniformly assess the costs of the fixed weight, bleed, resistance, etc. as a function of the fuel weight penalty, which is chosen as the optimization objective. The optimization variables consist of the compressor outlet temperature, cooling air flow rate, and fan duct heat exchanger structure size. The results show that the intermediate-stage bleed air and fan duct heat sink are more suitable when the Mach number is less than 2, but the ram air bleed is highly suitable for flight missions at a high Mach number. Especially at Mach 3.4–4.4, the ram air bleed mode can respond to the cooling and power demands with a simple architecture.

THE METHOD OF DETERMINING THE PUNCHING ACTION OF THE READY-MADE DAMAGE ELEMENT OF FRAGMENTATION OF HIGH-EXPLOSIVE WARHEAD

The subject of this article is the development of a method for determining the penetration capacity of ready-made elements of damage to a fragmentation and high-explosive warhead. The analysis of the existing methods of determining the penetration capacity of the damage elements indicates their inconsistency with the modern requirements for determining the power of means of destruction of combat parts of missiles, tactical and cruise missiles and aircraft bombs. Existing techniques that were proposed in the late 19th and early 20th centuries are based on the systematization of empirical research and are outdated and have a limited range of applications. Since they do not take into account the zone of contact interacting between the element of the impact and the obstacle and the critical angle of approach, which will determine the ricochet conditions. The article presents the results of the calculated determination of the penetration capacity of the ready-made damage element, taking into account the contact surface of the interaction of the damage element with the obstacle, which are obtained as a result of the calculation modelling in the Ansys Autodyn software complex. Ways to determine the penetration capacity of ready-made destruction elements of fragmentation and high-explosive warhead, taking into account their true geometric shape and excess kinetic energy are proposed in the article. The proposed technique can be used to create the penetration nomograms of the elements of the lesion. The application of computational modeling of the determination of nonograms of penetration ability will allow to abandon a large number of experimental studies with real undermining of the warhead, thereby reducing the cost and speeding up the process of creating, refining and adopting the newest samples of defeat equipment.

Aerodynamic configuration integration design of hypersonic cruise aircraft with inward-turning inlets

In this work, a novel airframe/propulsion integration design method of the wing-body configuration for hypersonic cruise aircraft is proposed, where the configuration is integrated with inward-turning inlets. With the help of this method, the major design concern of balancing the aerodynamic performance against the requirements for efficient propulsion can be well addressed. A novel geometric parametrically modelling method based on a combination of patched class and shape transition (CST) and COONs surface is proposed to represent the configuration, especially a complex configuration with an irregular inlet lip shape. The modelling method enlarges the design space of components on the premise of guaranteeing the configuration integrity via special constraints imposed on the interface across adjacent surfaces. A basic flow inside a cone shaped by a dual-inflection-point generatrix is optimized to generate the inward-turning inlet with improvements of both compression efficiency and flow uniformity. The performance improvement mechanism of this basic flow is the compression velocity variation induced by the variation of the generatrix slope along the flow path. At the design point, numerical simulation results show that the lift-to-drag ratio of the configuration is as high as 5.2 and the inlet works well with a high level of compression efficiency and flow uniformity. The design result also has a good performance on off-design conditions. The achievement of all the design targets turns out that the integration design method proposed in this paper is efficient and practical.

Effect of environmental radiation on the long-wave infrared signature of cruise aircraft

In the long-wave infrared radiation band (8–12 μm), the environmental radiation greatly affects the infrared signature of aircraft. The main aim of this paper is to assess the effects of atmospheric and ground radiance on the infrared signature of a cruise aircraft flying at high altitude. Firstly, the computation method of infrared signature is presented. The reverse Monte Carlo ray tracing method is applied to evaluate the effect of environmental radiation on the infrared signature of aircraft. The MODTRAN code is used to compute the atmospheric/ground radiance and atmospheric transmittance. Then, this method is validated with experimental data. Finally, the effect of environmental radiation on infrared signature of aircraft is discussed with the consideration of atmospheric condition, flight altitude, flight speed, and the emissivity of the airframe skin. The results show that the infrared signature of the lower surface of the aircraft is sensitive to the environmental radiation whose reflection contributes more than 20% to the total radiation at Mach 0.9. Increasing flight altitude, reducing atmosphere temperature and reducing flight speed will increase the ratio of reflected environmental infrared radiation intensity to total infrared radiation intensity of the aircraft. After considering the environmental radiation, the reduction amplitude of infrared signature of the aircraft by reducing the surface emissivity is decreased. The infrared signature of the aircraft lower surface is still significant compared to environment, even though the emissivity of the aircraft surface is reduced.

Installed performance evaluation of an air turbo-rocket expander engine

The propulsion plant of a prospective supersonic cruise aircraft consists of an air turbo-rocket expander and a dual-mode ramjet. A comprehensive numerical model was constructed to examine the performance of the air turbo-rocket during the supersonic acceleration of the vehicle. The numerical model comprised a one-dimensional representation of the fluid paths through the dual-mode ramjet, the air turbo-rocket combustor, the regenerator and the airframe-integrated nozzle, whereas the turbomachinery and the air turbo-rocket bypass were included as zero-dimensional models. The intake operation was based on the results of time-averaged Euler simulations. A preliminary engine analysis revealed that the installation effects restricted significantly the operational envelope, which was subsequently extended bypassing the air turbo-rocket. Hence the engine was throttled varying the mixture ratio and the fan compression ratio. Nevertheless, the performance was optimal when the demand from the air turbo-rocket matched the intake air flow capture. The heat recovery across the regenerator was found critical for the operation of the turbomachinery at low speed. The transition of the air turbo-rocket to ramjet operation was identified at Mach 4.5. During this regime, the propulsion plant was rather insensitive to the mixture ratio and was throttled with the air turbo-rocket throat area.

Supersonic Wing–Body Two-Level Wave Drag Optimization Using Extended Far-Field Composite-Element Methodology

The wave drag is an important factor that affects the supersonic flight efficiency and is closely related to the aircraft’s cross-sectional area distribution. A reasonable cross-sectional area distribution determines supersonic aircraft wave drag reduction. In this paper, an exploration of the supersonic aircraft wing–body shape parameterization using the class-shape-refinement-transformation technique and the method of extended far-field composite-element (EFCE) supersonic wing–body wave drag optimization are proposed. The method was applied to a supersonic civil transport wing–body shape optimization, yielding a 40.5% reduction of wave drag coefficient. In this way, the goal of reducing the wave drag in sufficient precision without any lift penalty was achieved. The EFCE wave drag optimization method is available in the conceptual design process of supersonic cruise aircraft to achieve high lift-to-drag ratio characteristics in supersonic flight conditions.

A Method for Aerodynamic Characteristic Analysis of Hypersonic Aircraft Based on Response Surface Model

Aerodynamic characteristic analysis of hypersonic cruise aircraft is more difficult than that of conventional aircraft, for the complex flow field simulation and inadequate amount of results under limited flight conditions. In this paper, numerical schemes applicable for hypersonic flow field are adopted to acquire a set of aerodynamic characteristics of a typical hypersonic cruise aircraft as sample data, based on which response surface models (RSM) are constructed to provide approximation of aerodynamic characteristics under any flight conditions within the design domain, finally the overall approximation performance of the response surface models are analyzed.