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The aircraft weight and balance problem (WBP) defined in terms of selecting unit load devices (ULDs) from an available set and allocating them to specified positions in cargo holds. It is critical to aircraft safety as well as operational efficiency and profit. It is an NP‐hard problem combining characteristics of the knapsack problem and generalized assignment problems, as well as complex practical constraints. A mixed integer programming model is constructed to maximize the total payload and minimize the center of gravity (CG) deviation by considering the constraints of positions, weights, balance etc. In contrast to prior studies using a constant delta to limit CG deviation, a new set of CG envelope constraints is proposed based on the variation in actual CG under different statuses and precise describing CG. A linearizing strategy is further designed to simplify the nonlinear CG envelope constraints. The proposed model can be solved using commercial optimization software within a few seconds and maintain the CG within the designated CG envelope in all test instances; in contrast, the other (CG deviation‐based) method may result in flight hazards in some scenarios.

The focus of this research was to understand the impact of batteries and electric motors on the weight and balance of the Parallel Electric-Gas Architecture with Synergistic Utilization Scheme (PEGASUS) concept. Because electrified aircraft propulsion components can comprise a large part of an aircraft’s weight, their integration has a significant impact on the center of gravity location and the related stability characteristics of the aircraft. We developed an analysis framework that enabled estimating the weight and volume of electrified aircraft components, their contribution to the center of gravity location, and the stability and performance of the aircraft. Trades were performed on battery placement, wing attachment point, and electric motor power to determine their impact on figures of merit such as maximum takeoff weight and block fuel. We identified batteries, and to a lesser extent electric motors, as having a large impact on aircraft center of gravity and the required horizontal tail size for longitudinal static stability. Placing electric component weight as far forward as possible resulted in reductions in the horizontal tail size required to maintain static stability. Configurations output from this framework were also evaluated using a six-degree-of-freedom simulation to quantify dynamic stability characteristics. Although shifting weight forward reduced horizontal tail size, it negatively impacted dynamic stability. This study confirms that mass property modeling and dynamic simulation, which are usually limited in conceptual design, are important for electrified aircraft concepts.

To maximize user experience in VR environments, optimizing the comfortability of head-mounted displays (HMDs) is essential. To date, few studies have investigated the fatigue induced by wearing commercially available HMDs. Here, we focus on the effects of HMD weight and balance on the physical load experienced by the user. We conducted an experiment in which participants completed a shooting game while wearing differently weighted and balanced HMDs. Afterwards, the participants completed questionnaires to assess levels of discomfort and fatigue. The results clarify that the weight of the HMD affects user fatigue, with the degree of fatigue varying depending on the center of mass position. Additionally, they suggest that the torque at the neck joint corresponds to the physical load imparted by the HMD. Therefore, our results provide valuable insights, demonstrating that, to improve HMD comfortability, it is necessary to consider both the balance and reduction of weight during HMD design.

A multi-scaled W&B demonstrator able of reproducing the same loading condition of an aircraft as a whole is designed, manufactured and experimentally characterized. Two laws of similarity guided the design of the demonstrator, concerning geometrical and weight aspects. Total weight and center of gravity position, including payloads are calculated starting from strain measures with Fiber Bragg Gratings (FBG) sensors installed on the landing gear system.

Air cargo transport is a growing industry in parallel with the growth in world trade and e-commerce. The global air cargo transport traffic getting busier, the importance of timely loading with minimum error is increasing. Therefore, digitalization of the air cargo loading process is needed. Assignment of Unit Load Devices (ULDs) to the specific positions on the freighter is performed by loadmasters by manual or semi-manual methods. This study aims to design a simulation model, which performs the ULD assignment automatically by simulating the loading process performed by the experienced loadmasters under the weight and balance constraints. The SEMMA (sample, explore, modify, model, assess) model is used while generating the simulation model. Fifty real-world loading orders were used to assess the performance of the model. The ULD assignment process by the simulation model and the loadmasters using semi-manual loading were compared with regard to time and center of gravity performance indicators. The results demonstrated that the simulation model can load all the given sets of ULDs, as efficiently as a loadmaster with a similar center of gravity in a shorter period of time. In conclusion, the proposed simulation model provides an efficient solution to the ULD assignment problem. However, the base model generated may be improved to address various real-world scenarios

Air cargo business is highly uncertain and dynamic. An airline needs to manage its air cargo loading operations with efficient cargo consolidation, accurate load planning, high level of safety precautions, limited cargo planning period and careful yield management in order to stay competitive and, in extreme situations, survive. To achieve these requirements for air freighters, a mathematical model and optimisation tool is developed to assist its air cargo load planning for the airline which handles about 327,500 tons of cargoes each year. A small improvement in its load planning, in terms of the best allocation and position of unit load devices to be stowed in an aircraft, can contribute substantially to its operational and financial performance. This study, applying in an cargo airline, aims to maximise cargo loading profit as well as improve the operation efficiency under the constraints of aircraft configuration, dangerous goods segregation, weight and balance, and flight safety. In future, the model can cater not only the cargo load planning of freighters but also passenger aircraft to improve the operational performance by considering over-sized cargoes.

Nowadays Unmanned Aircraft Vehicle (UAV) technology has been developed rapidly, including in Indonesia. LAPAN Surveillance UAV (LSU)-02 is one of the fixed-wing UAVs in Indonesia. It has been developed by LAPAN, the national aerospace research agency of Indonesia, for mapping and surveillance. The UAV has been equipped with automatic flight control system. However, the automatic system was not made based on the flight dynamic characteristic of the UAV. Therefore, optimization in certain flight condition becomes difficult. Hence, the aim of this research is to develop a mathematical model of LSU-02 for automatic control and guidance system design. A 6-DOF (degree of freedom) non-linear mathematical model was build based on aerodynamics, weight and balance, and propulsion of the UAV. The model is then evaluated by performing numerical simulation on the designed cruise condition. The simulation results show a dynamically stable characteristics for longitudinal motion and an unstable characteristic for lateral-directional motion.

Purpose This paper aims to present a knowledge-based fuel system, implementation and application, oriented towards its use in aircraft conceptual design. Design/methodology/approach Methodology and software tools oriented to knowledge-based engineering applications (MOKA) is used as a foundation for the implementation and integration of fuel systems. Findings Including fuel systems earlier in the design process creates an opportunity to optimize it and obtain better solutions by allocating suitable locations in an aircraft, thereby reflecting on the centre of gravity of the aircraft. Research limitations/implications All geometries are symbolic, representing a space allocation inside the aircraft for the fuel system. A realistic representation of the real components could be realized in detail design. Practical implications Fuel weight is a significant part of take-off weight and decisive in aircraft sizing and range estimations. The three-dimensional geometry provides a better estimation of the volume that is available to allocate the necessary entities. It also provides fast measures for weight and balance, fuel capacity, relative tank positions and a first estimation of piping length. Originality/value Fuel systems appear early in the design process, as they are involved in several first estimations. By using a knowledge-based engineering approach, several alternatives can be visualized and estimated in the conceptual design process. Furthermore, using the weights and centre of gravity at different angles of pitch and roll of each fuel tank, the aircraft could be optimized for handling qualities by using automatically generated system simulation models.

A novel approach to onboard in-flight weight and balance estimation systems is presented. Data from an Airbus A320 fleet from an airline were used to assess the feasibility of the approach. Simple flight mechanics in combination with statistics allowed for the identification of weight and center of gravity position using cruise angle of attack, Mach number and elevator deflection values. The good agreement between the theoretical model and the obtained values for the lift curve slope as a function of Mach as well as the standard error of the estimate for center of gravity position and cruise flying weight indicate that the method is sound. The major implication of this work is that the development of onboard and in-flight weight and balance systems can be significantly simpler than previous literature suggested. The impact of this paper could be immediate for airlines since all the tools required to implement the system as described are readily available. This could have an effect in operating costs, safety and environment.

This paper presents design of a typical Guided Flying Vehicle (GFV) using the multidisciplinary design optimization (MDO). The main objectives of this multi-disciplinary design are maximizing the payload’s weight as well as minimizing the miss distance. The main disciplines considered for this design include aerodynamics, dynamic, guidance, control, structure, weight and balance. This design of GFV is applied to three and six Degree of Freedom (DOF) to show comparison of simulation results. The hybrid scheme of optimization algorithm is based on Nelder-Mead Simplex optimization algorithm and Nondominated Sorting Genetic Algorithm II (NSGA II), called Simplex-NSGA II. This scheme is implemented for finding an optimal solution through the MDO. The Simplex-NSGA II method is a heuristic optimization algorithm that applies to multi-objective functions and the results are then compared with the most famous algorithms, like Nondominated Sorting Genetic Algorithm II (NSGA II) and Multi-Objective Particle Swarm Optimization (MOPSO). Simulation results demonstrate the superior performance of the Simplex-NSGA II over NSGA II and MOPSO. Also, it is used in this study in order to achieve an optimal solution using MDO in both 3DOF and 6DOF simulations of GFV to reach desirable performance index.

This paper provides information on the development of a landing gear Structural Health Monitoring (SHM) system that provides prognostic/diagnostic HUMS capabilities through direct load measurement in addition to strut servicing detection algorithms. The system provides advanced monitoring technology via the incorporation of new sensors integrated into the landing gear assembly. The direct load measurement approach is a paradigm shift from current methods of tracking fatigue damage of airframe landing gear systems and fuselage support structures, which depend on data collection of aircraft parameters recorded onboard at various sampling rates by SHM devices. The landing gear SHM provides direct loads measurement, weight/balance calculations, and the ability to perform Condition Based Maintenance (CBM) on the landing gear components. NAVAIR contracted with ES3 to support the development of the landing gear SHM via the Small Business Innovative Research (SBIR) program, via a Phase II award on the N121-043 topic. The proposed solution will be directly transferable to other Navy, military and commercial aircraft platforms. This paper will address the following topics in the area of HUMS and CBM: (1) advanced landing gear sensors for direct load measurement; (2) data fusion of direct loads monitoring data into fatigue life assessments; (3) paradigm shifts in aircraft maintenance utilizing strut servicing detection algorithms; (4) system verification and validation; and (5) safety and maintenance benefits. Prior work in the field of spectrum development and usage monitoring has typically focused on the aircraft structure, with assumptions translated to the landing gear components without any direct measurement. The benefits of usage monitoring can also be realized for landing gear. Direct loads measurement provides the ability to extend service life, remove components based on actual loading, improve safety, increase aircraft availability, and save maintenance costs with incorporation of CBM data into the maintenance practices. This paper advances the state-of-the-art via the miniaturization of sensors rated for the severe landing gear environment at a high Technological Readiness Level (TRL). The technological readiness of the landing gear SHM sensors has advanced, with several SHM sensors currently flying on aircraft for loads spectrum data collection purposes. This paper varies from prior publications for the European Workshop on Structural Health Monitoring, in that the prior applications concentrated on fluid level detection for CBM purposes—while this paper enhances the SHM system capability via the addition of direct load monitoring devices throughout the landing gear structural load path. Early development of the landing gear SHM system was presented at the 6th European Workshop on Structural Health Monitoring, in a paper regarding Aircraft Landing Gear Fluid Level and Landing Energy Monitoring System. The focus of that paper was the detection of improper fluid level and hard landings via the SHM system and sensors. This paper advances the prior publications via the addition of new direct load sensors into the SHM system, providing capabilities for fatigue damage tracking of landing gear components. The landing gear SHM system also allows data fusion of direct loads monitoring data into fatigue life assessments. This feature is provided via utilization of communication to platform HUMS and associated flight records for data assurance purposes. The interface and communication of the SHM control units to the aircraft HUMS equipment provides the ability to synchronize loads data, allowing for elimination/reduction of estimates on landing gear loads usage and service life. The SHM technology and algorithms also provide the ability for a paradigm shift in aircraft maintenance, utilizing strut servicing detection algorithms. Previously, automatic detection of an improperly serviced strut while the aircraft is on the ground was not possible. The incorporation of the SHM sensors and unique aircraft algorithms changes the maintenance approach, allowing for appropriate CBM based on the actual condition of the landing gear component service condition. High-fidelity laboratory testing of the SHM components using form, fit and functional hardware on a landing gear assembly has been completed. The completed demonstration testing included landing gear multi-body dynamic models, and validation of that model with the laboratory testing. The high-fidelity laboratory testing of loading events was accomplished on a full main landing gear assembly. The gear was instrumented as close as possible to the on aircraft configuration. The landing gear was secured in the test rig, in a similar manner to the landing gear attachment to the aircraft. The responses to loading events, as recorded by the SHM system, provided known loads, measured pressures and stresses to verify the landing gear hydrodynamic model. Landing gear model simulations were performed to create a virtual landing gear strut to predict the dynamic behavior of the strut servicing conditions, over a wide range of operational variables. Formal verification and validation of the landing gear model was accomplished, using a variety of available aircraft data—including instrumented flight test data. The SHM technology also provides a safety benefit with the improvement of weight and balance calculations. The landing gear sensors and associated control/interface units provide the ability to calculate actual weight and balance information. This information can then interface with other platform HUMS systems for use in improving maintenance practices, or enhancing crew safety during operations.

In this research, linear programming is the chosen method for optimizing aircraft center of gravity. The method was tested on the Airbus A320 and Boeing 757 aircraft. In order to successfully accommodate the special company requests, certain modifications are applied. Formulation of the proposed method enables the use of real data for the passenger and cabin baggage weight, instead of statistical values which are commonly used. The proposed method is simple to use, modify, upgrade or apply in existing systems.

Recently, VR technology has advanced rapidly, and many Head Mounted Displays (HMD) are sold by companies all over the world. In order to have good experiences in VR environment, comfortability of wearing HMD is essential. However, few studies investigated the feelings of fatigue on commercially available HMDs. In this research, we focused on the effects of weight and balance of HMD on physical load. We performed experiment with 188 participants using shooting game played under different weights and balances. After playing the game, the participants were asked to answer a questionnaire to assess the level of physical loads that they perceived. As the results, we clarified that the weight of HMD affects the feelings of fatigue, and its degree varies depending on the position of the center of gravity. The results suggest the joint torque at the neck can be an indicator of the physical load of HMD. Our results will be highly suggestive for designing new HMD.

NASA researchers, in a partnership with Boeing, are investigating a fuel-cell powered variant of the X-57 “Maxwell” Mod-II electric propulsion aircraft, which is itself derived from a stock Tecnam P2006T. The “Fostering Ultra-Efficient Low-Emitting Aviation Power” (FUELEAP) project will replace the X-57 power subsystem with a hybrid Solid-Oxide Fuel Cell (SOFC) system to increase the potential range of the electric-propulsion aircraft while dramatically improving efficiency and emissions over stock internal-combustion engines. Our FUELEAP safety analysis faces two primary challenges. First, the Part 23 certificated Tecnam P2006T is undergoing significant modifications to host the hybrid electric-propulsion system, and the challenge is to assure that the safety inherent in the stock aircraft (and subsequently in X-57 Mod-II) is not compromised by changes in avionics, aircraft structural loading, weight and balance, or other considerations. Secondly, because the SOFC power system has little (if any) relevant in-service precedent, our challenge is to assure that we identify and mitigate all reasonably plausible hazards introduced by unique FUELEAP equipage. We are investigating and utilizing Model-Based Safety Analysis (MBSA) methods to help us address these FUELEAP safety challenges. We captured aircraft-level system hazard conditions using instances of a SysML hazard block via aircraft-level Functional Hazard Analysis (FHA). Then, using SysML models of the FUELEAP architecture, we related the hazard conditions to initiating system events and possible mitigations, such as design architecture modifications or operational constraints. We are continuing to define our approach to MBSA by developing a component-by-component inventory of local failure modes and tracing their possible contribution to hazard conditions. Finally, we are applying an argument-based approach to FUELEAP assurance. Through a FUELEAP “safety case,” we are providing an explicit argument for FUELEAP safety by associating assurance evidence with overarching safety claims through a structured argument.

The premise is that in the future civil Air-to-Air Refuelling (AAR) will become an economic necessity if popular mass air travel is to continue. What is attempted is to provide a contemporary view of how such future operations could be safely undertaken. The intention is to predict the certification basis for demonstrating safe AAR operation of Cruiser-Feeder concepts. The necessary systems and aircraft functions are treated very much as they are today when civil certifying a large aeroplane type. The compliance demonstration required for environmental conditions, flight envelope, systems providing the necessary functionality, structural integrity, weight and balance are discussed. Applicable existing civil certification requirements are identified and where necessary expanded in scope to accommodate AAR operation. Where contemporary material does not supply appropriate guidance then corresponding safety criteria are proposed to address the deficiency. Lessons learnt from military AAR include the drive for interoperability. This has resulted in extensive efforts to standardise equipment and systems, which are equally applicable to civil AAR. Extremely useful advisory material exists, ranging from flight testing techniques to related safety. The importance of ensuring the consistency of failure condition categorisation at system and aircraft level is highlighted. The treatment of failures when two aircraft are in close proximity is something not considered by civil functional hazard analysis. The concept of AAR as an additional flight phase is introduced and affected system safety analyses identified. Examples of failure conditions that are not catastrophic at system level, but potentially could be at aircraft level during AAR are provided. Rendezvous scenarios are described to illustrate their influence on the certification basis. Combining such considerations with the factors that influence aircraft design leads to ramifications for handling qualities, performance and fuel system design. A viable and certifiable AAR configuration is consequently proposed. Consideration is given to treating operational certification in a progressive manner similar to existing LROPS (Long Range Operations).

Load master is authorized to be controlled all of loading stuffs for safety of flight such as passenger, baggage, cargo and e.t.c. There are many things are missed even though the weight and balance is the most important process. This study analyzes how the differences of C.G. by among ten load masters of each careers. This study is tested how load-master takes load-control by the respective result based on differences of each practical experiences, gender and a number of certification. In result, all of load masters set C.G. on the stability region. But the practical experience of load master is relative to set better C.G. for economical effectiveness of weight and balance control work.

Blended Wing Body concept offers several advantages compared to traditional aircraft tube and wing concept. The advantages mainly come from the distributed aerodynamic and structural loads, which leads to better aerodynamic performance as well as lighter structural weight. Most of the existing studies were focused on big transport aircraft carrying 400 to 800 passenger. In this study, a conceptual design of a business jet aircraft applying blended wing body concept is carried out. The market forecast for this category of aircraft can reach up to 24000 aircraft in the next 20 years. The possibility of having larger cross section is one of the competitive advantages, notably in a long-range flight. The requirement stated was to fly a trans-atlantic flight and carrying up to 18 passengers. It has to have a low floor height permitting easy passenger access. The design process consists of initial weight estimation, initial sizing, and preliminary aerodynamics, weight and balance and performance analysis. Some design consideration specially related with Blended Wing Body concept will be discussed, i.e. take-off and landing aerodynamics, structural concept, stability and control. The final design resulted in twin-engine aircraft, a Maximum Take-off Weight of 44 ton, cabin floor area of 6.9 x 10 m, winglet and split rudder for directional stability and control, and elevons for longitudinal and lateral control.

This chapter describes the most significant regulatory involvement a developer is likely to encounter in the certification of avionics. It discusses certification activities appropriate to the US Federal Aviation Administration (FAA). The chapter discusses some representative differences between FAA practices and those of Europe’s Joint Aviation Authorities. The Federal Aviation Regulations provides several different forms of approval for electronic devices installed aboard civil aircraft. Of these, most readers will be concerned primarily with approvals under the technical standard order system, approvals under a supplemental type certificate (STC), or approvals as part of a type certificate, amended type certificate, or service bulletin. The STC process assumes modification of at least one prototype aircraft. It is in the aircraft modification that all the engineering analysis—aircraft performance, structural and electrical loading, weight and balance, human factors, and so on—comes together. New avionics equipment installed as part of an STC will often lead to new and different procedures for flight crews.

The maximum horizontal impact load experienced by the aircraft landing gear wheel during landing is an important parameter in landing gear safety design and performance analysis. Using a non-contact photoelectric testing method, this paper experimentally obtained the dynamic data of the instantaneous rotational speed for an aircraft wheel when it contacts the platform. Moreover, the kinetic relationship between the transient rotational speed of the aircraft wheel and the force of horizontal impact was established. According to the change data measured from the transient rotational speed of the aircraft wheel, the maximum horizontal impact load at the instant of aircraft contact with the platform was calculated. To verify the accuracy of this method, in which changes in rotational speed determine the horizontal impact load, verification equipment without lateral constrains was established to test the horizontal impact load of the aircraft wheel. The experimental results showed that the horizontal impact load determined by the transient rotational speed of aircraft wheel is consistent with the results obtained by self-built measurement equipment without lateral constrains.

A multiphase approach is considered for the design of an aircraft anti-skid braking control system. A four-phase scheme, based on angular acceleration thresholds and estimated slip rate values, is used to stabilize the aircraft system around an optimal braking condition. The change in vertical load due to variable lift force and the ensuing variable tire deflection are considered in the verification models. A hybrid controller is proposed, designed to achieve a stable wheel slip limit cycle. The simulations confirm that the proposed control scheme is a valid solution for the anti-skid of an aircraft main Landing Gear (LG). The actuator time delay and dynamic bandwidth influence on the aircraft stop distance has been studied.

The aim of this study was to implement an optimal arrangement of equipment, instrumentation and medical personnel based on the weight and balance of the aircraft and to transfer humanitarian aid in a drone, by implementing artificial intelligence algorithms. This is due to the problems presented by the geographical layout of human settlements in southeast of the state of Chihuahua. The importance of this research is to understand the multivariable optimization associated with the path of a group of airplanes associated with different kinds of aerial in order to improve the evaluation of flooding and to send medical support and goods; to determine the optimal flight route, including speed, storage and travel resources. To determine the cost–benefit, this has been partnered with a travel plan to rescue people, which has as its principal basis the orography airstrip restriction, although this problem has been studied on several occasions by the literature failed to establish by supporting ubiquitous computing for interacting with the various values associated with the achievement of the group of drones and their cost–benefit of each issue of the company and comparing their individual trips for the rest of group. There are several factors that can influence in the achievement of a group of drones for our research. We propose the use of a bioinspired algorithm.

There exist vertical long ropes that need periodical inspection for infrastructure maintenance. Development of lifting devices along with the ropes is demanded for steady observation at human-unreachable positions. The lifting device based on alternating rotation hoist has been proposed for carrying inspection devices to the target level. This research realizes an automatic movement of the lifting device by computer control for the improvement of speed and accuracy. Furthermore, a new frame which is equipped with the inspection device is developed for saving weight and keeping balance. Through experiments, performances of the new frame and control system are evaluated from the aspects of speed, accuracy, weight and balance. The results show that the newly developed frame has advantages in weight and speed. The inclination of frame and ascender is measured. The error values are calculated to analyze the accuracy. And the developed control system can control frame effectively.

This paper presents a new algorithm that predicts the quantity of fuel burned by an aircraft flying at a constant speed and altitude. It considers the continuous fuel burn rate variation with time caused by the gross weight (and centre of gravity position) modification due to the fuel burn process itself. The algorithm was developed for use by the Flight Management System (FMS) and employs the same aircraft performance data as the existing FMS fuel burn prediction algorithms. The new fuel burn method was developed for aircraft models that use the centre of gravity position as well as for models that do not consider the centre of gravity position. This algorithm was developed for normal flight conditions. Algorithm performances were evaluated for two aircraft models: one for models that use an aircraft’s centre of gravity position – a more complex and computing intensive method, and one for those that do not use the centre of gravity position. The validation data were generated based on the information produced on a CMC Electronics – Esterline FMS platform that used identical aircraft models and performance data for identical flight conditions.

Aiming at the problem that the flight performance of multi-rotor with high load ratio is easily affected by the change of the center of gravity of load, the variable center of gravity model of multi-rotor was established, and the influence of the center of gravity on the aircraft was analyzed, and on the basis of this, the solution was put forward. Finally, the effects of the center of gravity change on flight performance and the effectiveness of the solution were verified through MATLAB simulation experiments.

Aircraft center of gravity control is an important technology to gain potential performance benefits. In-flight center of gravity control is usually based on a fuel transfer system which enables fuel to be transferred among fuselage fuel tanks and then adjust aircraft center of gravity. The aim of this article is to systematically investigate the influences of partial fuel pump failures on the CG control performance for the high-speed aircraft. As the basis, modeling process of center of gravity varying with fuel transfer is put forward and the sliding model based control law for center of gravity is presented in detail. Then the detailed simulation analysis of possible failures effects on center of gravity tracking performance for the high-speed aircraft is carried out, including different one or two fuel pump failures. Finally, the corresponding emergency strategies for the aircraft center of gravity control system are proposed.

Set up a kinetic analysis model to study ground-maneuver-stability of multiple-bogie undercarriage. The calculations and simulations demonstrate that the allowable nose wheel steering angle and aircraft's ground-maneuver stability could be improved by decreasing height of C.G of aircraft, increasing aircraft's weight, decreasing aircraft's velocity and decreasing engine thrust on condition that aircraft's configuration is identical and aircraft's safety and performance do not reduce.

A sliding mode based active center of gravity control law is designed for the aircraft with center of gravity varying during flight. As the basis, the general scheme of the active center of gravity control system is proposed, and the mathematic model of aircraft center of gravity varying with fuel transfer is established. Then, the center of gravity control law based on sliding mode control theory is designed, which can guarantee the control error converges to zero in finite time. Finally, simulations are carried out to verify the proposed method. Simulation results show that the sliding mode based center of gravity control law can accurately and rapidly track the given command for various different fuel transfer conditions.

This paper presents a safety risk management model based on possibility theory to adapt and develop the Safety Management System (SMS) in airline operation control. Fuzzy sets are used to represent the key elements among crew, meteorology, weight and balance, dispatch, coordinate, and systematic operation control by risk indicators. A fuzzy integrated judgment model is also proposed, which would quantity both the possibility and necessity of the safety space. The implications of this strategy for practitioners and future work are discussed.

The goal of this paper is the development of a new mixed integer linear program designed for optimally loading a set of containers and pallets into a compartmentalised cargo aircraft. It is based on real-world problems submitted by a professional partner. This model takes into account strict technical and safety constraints. In addition to the standard goal of optimally positioning the centre of gravity, we also propose a new approach based on the moment of inertia. This double goal implies an increase in aircraft efficiency and a decrease in fuel consumption. Cargo loading generally remains a manual, or at best a computer-assisted, and time-consuming task. A fully automatic software was developed to quickly compute optimal solutions. Experimental results show that our approach achieves better solutions than manual planning, within only a few seconds.

Simple, approximate relations are developed that allow rapid, yet accurate, estimates of the principal mass characteristics of modern civil transport aircraft. The method is based upon an ‘aircraft mass hypothesis’ combined with data taken from Type Certificates and manufacturers’ information for 44 different Boeing and Airbus aircraft types and variants. This hypothesis links the masses of the various aircraft components, including the operational items, and the maximum payload mass to just two parameters; the passenger cabin floor area, as exemplified by the ‘design passenger number’ and the ratio of maximum zero fuel mass to maximum take-off mass, as exemplified by the non-dimensional ‘design range’. In this context, the ‘design range’ is defined as the maximum possible range when the aircraft is operating at its maximum zero fuel mass. This approach leads to the formulation of three basic ‘laws’ for aircraft mass. Special consideration is given to the identification and separation of those components that are fixed either by the design, or the certification process, and those that are free, within defined limits, to be chosen by the operator.

The Aircraft Weight and Balance Problem (AWBP) is a real-world combinatorial optimisation problem. From a set of available cargo units, the most profitable selection has to be loaded into an aircraft, while satisfying a large number of safety constraints, of which some are non-linear. In this paper a mixed integer programming approach to the AWBP is introduced. The aim is to increase loaded cargo value and to decrease deviation of the optimal centre of gravity. Experimental results on real life data show significant improvements on both objectives compared to the results obtained by an experienced planner.

Conservative flight performance of an aircraft with constant load factor was analysed by Miele. In this paper, conservative flight with a varying load factor is considered. The load factor being a control parameter is varied in such a way that it gives rise to closed-form solutions to aircraft motion. With this typical load factor function the geometry of the trajectory has been found to be a loop as is with Miele’s analysis. Finally, the range and endurance have been determined for two different constant parameters associated with the load factor function.

determined by the maximal lateral loading conditions identified in published studies of instrumented in-service passenger aircraft. The performance of the turn can be assessed over the entire operational range in terms of the actual loads experienced at individual landing gears. Recent studies by the Federal Aviation Administration of instrumented aircraft have been limited to investigating the lateral loads experienced at the aircraft’s center-ofgravity position. The results presented here show that this information is insufficient to predict the actual loads experienced by individual landing gears, especially for the nose gear, which is found to experience considerably higher lateral loads than predicted by the corresponding loads at center of gravity. These findings are shown to be robust with respect to changes in the aircraft’s mass and the criterion used to define the limits of the operating regions.

The payload capability of many aircraft is typically limited by the maximum takeoff weight. However, when in flight, many aircraft can handle larger loads, but this capability is rarely exploited. The possibility of using an aircraft cable system to lift large payloads is studied in this paper. This concept has been studied previously for single-aircraft, single-cable systems in some detail. The major limitation of such systems is the need to fly very tight circles at high-lift coefficients. This limits the size of the payload that can be retrieved with such a system. To overcome this, it is possible to employ a multiple-aircraft, multiple-cable system to allow greater control of the payload position as well as allowing a much wider choice of aircraft speed and lift coefficient. In this paper, an object-oriented cable modeling approach is used to study the relative equilibria of such a system. Parametric studies of different payload masses and aircraft orbit parameters are presented, and the dynamic stability is analyzed. Collision-free transition trajectories that take the multiple-aircraft system from circular flight into straight and level flight are derived.

The ability to compute the weight and balance of a helicopter in flight under general conditions is an enabling technology for future condition based maintenance systems as well as advanced automatic flight control systems. This paper creates a real-time weight and balance estimation algorithm using an extended Kalman filter framework. To highlight estimation characteristics, the algorithm is exercised on the OH-6A helicopter in a variety of flight regimes. The algorithm is examined in hover and forward flight as well as situations where loads are dropped and picked up in flight. For sample scenarios considered, the algorithm quickly estimates station line and butt line mass center position (1 s) and more slowly converges on helicopter weight (10 s). To estimate the helicopter waterline, the algorithm requires modest flight maneuvering where a nonzero roll rate is present. The algorithm is also shown to be reasonably robust relative to sensor and model errors with increasing estimation error with increasing levels of model mismatch and sensor error.

The Brazilian F-5 was submitted to avionics and weapons upgrade. This �new � aircraft has proven to be heavier and more capable. A comprehensive damage tolerance analysis is being performed to evaluate how the new mission profiles and weight distribution may affect the airframe structural integrity. Operational data were collected at the Brazilian Air Force Bases where the fighter is flown. Software was developed in order to acquire, filter and analyze flight data. This data was used for comparison between the pre and post modernization mission profiles and to determine the stress level in each of the known aircraft fatigue critical locations (FCL). The results show that the change in aircraft weight and balance and the new operational profile can significantly change the inspection intervals of certain fatigue critical locations of the structure. A preliminary result for the horizontal tail has shown that this component will have a much more restrictive maintenance schedule to assure flight safety.

One of the crucial challenges of the aviation industry in upcoming years is to reduce emissions not only in the vicinity of airfields but also in cruise. Amongst other transport methods, airplanes emissions count for 3% of the CO 2 emissions. Initiatives to reduce this include not only investing in more fuel-efficient aircrafts or adapting existing ones to make them more efficient (e.g. by fitting fuel-saving winglets), but also more actively researching novel propulsion systems that incorporate environmentally friendly technologies. The Boeing Company through its European subsidiary, Boeing Research and Technology Europe (BR&TE) in collaboration with industry partners throughout Europe is working towards this goal by studying the possible application of advanced batteries and fuel-cell systems in aeronautical applications. One example is the development of a small manned two-seater prototype airplane powered only by proton exchange membrane (PEM) fuel-cell stacks, which runs on compressed hydrogen gas as fuel and pressurized air as oxidant, and Li-ion batteries. The efficient all composite motorglider is an all electric prototype airplane which does not produce any of the noxious engine exhaust by-products, such as carbon dioxide, carbon monoxide or NO x , that can contribute to climate change and adversely affect local air quality. Water and heat are the only exhaust products. The main objective is to demonstrate for the first time in aviation history a straight level manned flight with fuel-cells as the only power source. For this purpose, the original engine of a super Dimona HK36TTC glider from Diamond Aircraft Industries (Austria) was replaced by a hybrid power system, which feeds a brushless dc electrical motor that rotates a variable pitch propeller. Amongst the many technical challenges encountered when developing this test platform are maintaining the weight and balance of the aircraft, designing the thermal management system and the power management between the two power sources [N. Lapena-Rey, J. Mosquera, E. Bataller, F. Orti, SAE 2007 Aerotech Congress & Exhibition, 2007 (Publication number: 2007-01-3906)]. The demonstrator airplane constitutes an example of the successful implementation of novel clean power sources in aviation. The detailed description of the airplane and its subsystems is given elsewhere [N. Lapena-Rey, J. Mosquera, E. Bataller, F. Orti, SAE 2007 Aerotech Congress & Exhibition, 2007 (Publication number: 2007-01-3906)]. This paper focuses specially on the power sources design and pre-flight tests giving special attention to those requirements derived from aerospace applications.

We propose two nonlinear approaches for the autonomous transition control of two vertical/short takeoff and landing aircraft. The first aircraft, referred to as the V-1 aircraft, is a fixed-wing fighter equipped with a vectored thrust and lift fan. The second aircraft is a tilt rotor. We provide nonlinear models for the transition from hover to forward flight. Both control problems are nonlinear, nonaffine, and have redundant inputs. Furthermore, some control inputs must reach specified end values when the transition has been completed. However, there is a key difference between the two aircraft. The V-1 aircraft can be controlled to rapidly increase velocity and pitch angle simultaneously, whereas the tilt rotor aircraft cannot, because linear and angular acceleration require conflicting control inputs from hover. For the V-1 aircraft, a method is applied whose key point is that a new controlled variable is appended to decrease the redundancy of the control inputs. For the tilt rotor aircraft, a second method is applied that transfers the problem into a two-time-scale problem. Nonlinear optimization is adopted to make the specified input approach the specified end value. We also consider problems associated with cases in which insufficient control power limits the ability of the aircraft to achieve commanded controlled variable rates. Simulations show that both approaches achieve the transition control successfully.

This article describes a WBS (Weight and Balance System) for a large aircraft of the pressure distortion compensation device. This system can eliminate the pressure distortion of hydraulic buffer due to the friction, thus to improve the accuracy of real-time measurement for overall weight of aircraft. In this article, which introduce the principle and structure of the aircraft real-time measurement of overall weight in WBS, and due to friction is present in hydraulic buffer, analysis the cause of tolerance of weight measurement, then correct it by using pressure distortion compensation device. After that, the structure of the device is introduced, and describes schemes of aircraft equipped with device, and a step for correction of weight measurement. Meanwhile, it shows the compensation effect chart of weight real-time measurement. Technical and accuracy parameters which measurement overall weight is comparison before and after by using the device, to evaluate the significance and the possibility of implement on the aircraft.

This paper describes the design and conduct of the strain-gage load calibration ground test of the SubsoniC Research Aircraft Testbed, Gulfstream III aircraft, and the subsequent data analysis and results. The goal of this effort was to create and validate multi-gage load equations for shear force, bending moment, and torque for two wing measurement stations. For some of the testing the aircraft was supported by three airbags in order to isolate the wing structure from extraneous load inputs through the main landing gear. Thirty-two strain gage bridges were installed on the left wing. Hydraulic loads were applied to the wing lower surface through a total of 16 load zones. Some dead-weight load cases were applied to the upper wing surface using shot bags. Maximum applied loads reached 54,000 lb. Twenty-six load cases were applied with the aircraft resting on its landing gear, and 16 load cases were performed with the aircraft supported by the nose gear and three airbags around the center of gravity. Maximum wing tip deflection reached 17 inches. An assortment of 2, 3, 4, and 5 strain-gage load equations were derived and evaluated against independent check cases. The better load equations had root mean square errors less than 1 percent. Test techniques and lessons learned are discussed.

Modern versions of supersonic jet fighter aircraft have several different weapon loading configuration to support air-to-air combat and air-to-ground delivery of weapon modes. These various aircraft loading conditions could result in asymmetric configurations to the aircraft once delivered. These asymmetric configurations could result in decreased handling qualities for the pilot maneuvering, stability, control issues and aerodynamic performance of the aircraft. In order to eliminate or decrease these adverse impacts on the pilot`s ability, development of T-50 flight control law attempts to control the aircraft in both longitudinal and lateral-directional axes. Especially, the design of the lateral-directional roll axis control laws, utilizing a simple roll rate feedback structure and gains such as F-16, is developed for the T-50 aircraft to meet the aircraft`s design requirements. Consequently, it is found that the improved control law decreases the roll-off phenomenon in lateral axes during pitch maneuver.

This paper shows how the dynamics of a rotor in a maneuvering aircraft changes according to the operation of the aircraft. The mathematical model of an unbalanced rotor system located in the maneuvering aircraft is derived. The dynamic characteristics of the rotor running at a constant angular speed or a constant acceleration are studied under the assumptions that the aircraft maneuvers only in a vertical plane and that the pitching angle and the flight path inclination of the aircraft are equal. The effects of gravity and unbalance parameter are considered. The results show that the unbalanced response of a rotor in an aircraft is obviously influenced by the aircraft’s flying status.

In a simultaneous paired approach to closely-spaced parallel runways, a pair of aircraft flies in close proximity on parallel approach paths. The longitudinal separation between the aircraft must be maintained within a range that avoids wake encounters and, if one of the aircraft blunders, avoids collision. To increase operational availability, the approach procedure must accommodate a mixture of aircraft sizes and, consequently, approach speeds. In these procedures, the slower aircraft is placed in the lead position. The faster aircraft maintains separation from the slow aircraft in a dependent operation until final approach and flies independently afterward. Due to the higher approach speed of the fast aircraft, longitudinal separation will decrease during final approach. Therefore, the fast aircraft must position itself before the final approach so that it will remain within the safe range of separation as separation decreases. Given the approach geometry and speed schedule for each aircraft, one can use kinematics to estimate the separation loss between a pair of aircraft. A kinematic model can complement fast-time Monte-Carlo simulations of the approach by enabling a tailored reduction in the variation of starting position for the fast aircraft. One could also implement the kinematic model in ground-based or on-board decision support tools to compute the optimal initial separation for a given pair of aircraft. To better match the auto-coupled flight of real aircraft, the paper derives a kinematic model where the speed schedule is flown using equivalent airspeed. The predicted time of flight using the equivalent airspeed kinematic model compares well against a high-fidelity aircraft simulation performing the same approach. This model also demonstrates a modest increase in the predicted loss of separation when contrasted against a kinematic model that assumes the scheduled speed is true airspeed.

Having determined the size and placement of all the major components of the aircraft in the previous chapters, a refined weight estimate is carried out and the center of gravity location is determined. Techniques for refining the weight estimate are presented. The limit load factor and the design dive speed requirements are outlined. Methods are presented for making refined estimates of the weight of the various aircraft components: wing group, fuselage group, landing gear group, tail group, propulsion group, nacelle group, flight controls group, auxiliary power unit group, instrument group, hydraulic and pneumatic group, electrical group, avionics group, equipment and furnishing group, air conditioning and anti-icing group. The concept of the “all-electric” airplane is discussed. The center of gravity of the aircraft is estimated by considering the contribution of the weight and center of gravity of the major components: wing, fuselage, landing gear, tail, and propulsion groups.

The present contribution introduces a mixed integer linear programming model as a decision support tool for air cargo load planning. The main objective for the model is to find the most profitable selection from a set of cargo to be loaded on an aircraft. The secondary objective is to minimize the deviation between the aircraft’s centre of gravity, and a known target value so as to reduce fuel consumption and improve stability. The model is subject to a large number of constraints that ensure structural integrity and stability of the aircraft, as well as the safety of the cargo and crew. A set of additional constraints guarantees safe and efficient loading and unloading. Experimental results on real-life data show that the model outperforms human expert planners on both objectives, while remaining computationally fast enough for interactive use. This advocates the use of such a decision support model for all air cargo load planning. publisher: Elsevier articletitle: Automatic air cargo selection and weight balancing: A mixed integer programming approach journaltitle: Transportation Research Part E: Logistics and Transportation Review articlelink: http://dx.doi.org/10.1016/j.tre.2013.12.013 content_type: article copyright: Copyright © 2014 Elsevier Ltd. All rights reserved. ispartof: Transportation Research E, Logistics and Transportation Review vol:65 issue:1 pages:70-83 status: published

This paper presents an evaluation of available aircraft types and demonstrates the lack of a low cost aircraft optimized for fighting forest fires, spreading viscous liquids for land reclamation and spraying pesticides over forest areas. It shows that among critical areas demanding special consideration are (a) the development of mathematical models for the water bomb-aircraft separation and the aircraft transient dynamics following separation, (b) the identification of parameters influencing the coherence of the water column and effectiveness of water delivery for fire fighting, (c) choice of aircraft configuration and (d) hopper configuration. Design and numerical analysis have led to the selection of the biplane as the best aircraft. The main aerodynamic characteristics for the selected aircraft have been computed by means of panel methods, the so-called modified Hess method for thick wings and bodies and/or the vortex lattice method for thin lifting surfaces. Different gaps and staggers and their influence on aerodynamic characteristics have been analysed. It has been found that shifting the lower wing rearwards (positive stagger) while keeping the angle of attack constant results in a small increase of induced drag and an almost constant value of the lift curve slope as well as an increase in the pitching moment curve slope. The increase of drag is disadvantageous, whereas the increase of the pitching moment curve slope means that the neutral point of stability is moved forward (a disadvantage from the stability point of view). The influence of the biplane configuration on downwash in the vicinity of the horizontal plane and aircraft dynamic stability is also discussed. Another important concept—developed at PZL-Okecie and presented in this paper—consists in using parts from existing aircraft. The pilot's cabin, the rear part of the fuselage with control surfaces and wings originate from the PZL-106 ‘Kruk’. This diminishes the cost of design and prototype construction as well as of the cost of aircraft production. It has been shown that an important cost factor in the operation of such a fire-fighting aircraft is the weight of the agent which may be carried for the same fuel consumption. This cost factor, representing the economical efficiency of a fire-fighting aircraft, has been computed and compared for a number of fire-fighting aircraft. The design under consideration (called the PZL-240 ‘Pelikan’) has the above-mentioned factor equal to 14, whereas the average value for other aircraft is about 8.

We introduce a simple blending method for parametric curves and surfaces that produces families of parametrically defined, G n –continuous blending curves and surfaces. The method depends essentially on the parameterizations of the curves/surfaces to be blended. Hence, the flexibility of the method relies on the existence of suitable parameter transformations of the given curves/surfaces. The feasibility of the blending method is shown by several examples. The shape of the blend curve/surface can be changed in a predictable way with the aid of two design parameters (thumb weight and balance).

Sufficient conditions of technical stability in measure of dynamic states of an aircraft running on a runway with a high speed are obtained. It is assumed that the weight of the aircraft exceeds slightly the wing lift, the crosswind is strong, and the initial distributions of process parameters are arbitrary. The dependence between the eigenvalues of the positive definite quadratic form constructed and the criteria of technical stability is formulated