International Standard Atmosphere Research Articles

Variations of the zero degrees isotherm and environmental lapse rate recorded with ADSB and Mode S EHS messages

This study focuses on analyzing the altitude of the zero degrees isotherm and variations in the environmental lapse rate (ELR) using the publicly available data collected from Automatic Dependent Surveillance–Broadcast (ADS-B) and Mode S Enhanced Surveillance (EHS) messages emitted by airplanes over a five-month period in 2021. The data was gathered using a professional receiver stationed in Bucharest (Romania). The aviation messages were decoded and the air temperature and pressure were determined, at the location of the airplane. The method has the advantage of the continuous messages that are emitted by the aircrafts during flight that allow instantaneous determination of the meteorological parameters at no additional costs. It can also be extended to permit almost real time maps of the ELR. When data was analyzed and a standard ELR value of 6.5 K/km is employed it was observed that the mean altitude of the 0 degrees isotherm exhibits a seasonal increase during the summer months, with an average altitude of 2874.2 m. The highest recorded altitude of the 0 degrees isotherm was found to be 5346.8 m, near Alexandria city (Romania), on 22.07.2021. Using a standard Least Mean Square algorithm alongside the International Standard Atmosphere pressure formula, the ELR values were calculated from pressure measurements data. The resulting mean ELR for the five-month period was determined to be 5.1331 K/km, slightly lower than the standard value.

Investigation of Onboard Power Generation Method Using Waste Energy at Altitude Conditions

Abstract Demand for increased onboard power generation on small aircraft is ever growing due to increased electrical power demand. The objective of this study is to investigate onboard power generation using waste energy at various altitude and ambient conditions. Experiments were performed on a 2-liter turbocharged direct-injection intermittent combustion engine fueled with jet fuel (F-24). The turbocharger has an integrated motor-generator system that is designed to generate electricity using the enthalpy in the engine exhaust, or to use as a motor to rotate the turbocharger using the energy stored in onboard energy storage device. An electrically actuated wastegate was used to control useful exhaust flowrate that is used to generate electricity. An external power supply system was used to emulate the energy storage. The engine was installed in the US DEVCOM Army Research Laboratory?s altitude chamber in which absolute outside air pressure was controlled up to 37.6 kPa to simulate the altitude conditions up to 7,620 m (25,000 ft). Three different ambient conditions were selected to represent International Standard Atmosphere (ISA), polar, and tropical conditions. Based on the experimental results, a response surface model was developed to predict the power generation of the motor-generator integrated turbocharger using waste energy as a function of altitude, engine power, and the wastegate position. The result from this study provides an insight into onboard power generation method using waste energy at altitude conditions.

On the conditions for absolute minimum fuel burn for turbofan powered, civil transport aircraft and a simple model for wave drag

Abstract In a recent series of papers, Poll and Schumann have been developing a simple model for estimating fuel burn for turbofan powered, civil transport aircraft for a given mass, Mach number and flight level and in a specified ambient temperature profile for all phases of flight. This paper focuses upon the combination of Mach number and flight level at which an aircraft cruises with the absolute minimum fuel burn. For a given aircraft type, the information necessary to determine these conditions must be specified and this poses a challenge. An initial attempt to obtain these data has been described previously by the first author. In this paper, the optimum conditions are found using a completely different approach. Starting from first principles and using established theory, the equations governing the situation where engine overall efficiency and airframe lift-to-drag ratio both have local maxima at the same flight condition are developed. This special case is termed the “design optimum” condition and, for a specified aircraft mass and a specified atmospheric temperature versus pressure profile, it gives the lowest possible fuel burn for any aircraft and engine combination. The design optimum occurs at a particular Mach number and Reynolds number, and it is a fixed characteristic of the aircraft. The analysis reveals the significance of Reynolds number variations, wave drag, including its derivatives with respect to both lift coefficient and Mach number, and the atmospheric properties. Whilst wave drag is notoriously difficult to determine accurately, it is found that solutions to the equations are not particularly sensitive to the accuracy of this quantity. Consequently, a simple, physically realistic model can give good results. An appropriate model is developed and a complete, approximate solution is obtained. Taking the International Standard Atmosphere as the design atmosphere, results are presented for the 53 aircraft types previously considered by Poll and Schumann. Relative to the design optimum conditions, when Reynolds number is constant and wave drag is zero, compressibility alone reduces L/D by about 5%, reduces lift coefficient by about 1.5% and increases drag coefficient by about 3.5%. Reynolds number variation has little effect upon L/D, but it reduces lift coefficient and drag coefficient by a further 7% and 8% respectively. The reduction in lift coefficient has a significant impact on the optimum cruise flight level. In general, an aircraft’s operating optimum will not coincide with its design optimum, but deviations are expected to be small. Therefore, using the design optimum solution as a reference point, an improved version of the operating optimum estimation method described by Poll and Schumann in previous work is developed. This allows the estimation of the conditions for absolute minimum fuel burn for an aircraft of given mass flying thorough any atmosphere. Updated coefficients for the 53 aircraft types are given.

Estimation of the Standard Atmospheric Earth Model Parameters at 86 km Altitude

Utilizing the Turbo C programming language, the atmospheric earth model is created from sea level to 86 km. This model has been used to determine atmospheric Earth parameters in this study. Analytical derivations of these parameters are made using the balancing forces theory and the hydrostatic equation. The effects of altitude on density, pressure, temperature, gravitational acceleration, sound speed, scale height, and molecular weight are examined. The mass of the atmosphere is equal to about 50% between sea level and 5.5 km. g is equal to 9.65 m/s2 at 50 km altitude, which is 9% lower than 9.8 m/s2 at sea level. However, at 86 km altitude, g is close to 9.51 m/s2, which is close to 15% smaller than 9.8 m/s2. These results have been compared with an international standard atmosphere. The presumed atmosphere model differs significantly from the actual atmosphere because weather fluctuations are not taken into consideration in this model.

Trend in tropopause warming and its influence aircraft performance

PurposeTemperature anomalies in the upper troposphere have become a reality as a result of global warming, which has a noticeable impact on aircraft performance. The purpose of this study is to investigate the total air temperature (TAT) anomaly observed during the cruise level and its impact on engine parameter variations.Design/methodology/approachEmpirical methodology is used in this study, and it is based on measurements and observations of anomalous phenomena on the tropopause. The primary data were taken from the Boeing 747-8F's enhanced flight data recorder, which refers to the quantitative method, while the qualitative method is based on a literature review and interviews. The GEnx Integrated Vehicle Health Management system was used for the study's evaluation of engine performance to support the complete range of operational priorities throughout the entire engine lifecycle.FindingsThe study's findings indicate that TAT and SAT anomalies, which occur between 270- and 320-feet flight level, have a substantial impact on aircraft performance at cruise altitude and, as a result, on engine parameters, specifically an increase in fuel consumption and engine exhaust gas temperature values. The TAT and Ram Rise anomalies were the focus of the atmospheric deviations, which were assessed as major departures from the International Civil Aviation Organizations–defined International Standard Atmosphere, which is obvious on a positive tendency and so goes against the norms.Research limitations/implicationsNecessary fixed flight parameters gathered from the aircraft's enhanced airborne flight recorder (EAFR) via Aeronautical Radio Incorporated (ARINC) 664 Part 7 at a certain velocity and altitude interfacing with the diagnostic program direct parameter display (DPD), allow for analysis of aircraft performance in a real-time frame. Thus, processed data transmits to the ground maintenance infrastructure for future evaluation and for proper maintenance solutions.Originality/valueA real-time analysis of aircraft performance is possible using the diagnostic program DPD in conjunction with necessary fixed flight parameters obtained from the aircraft's EAFR via ARINC 664 Part 7 at a specific speed and altitude. Thus, processed data is transmitted to the ground infrastructure for maintenance to be evaluated in the future and to find the best maintenance fixes.

International standard atmosphere - a tool for technological measurement sovereignty in the aerospace industry

When designing aircraft for agriculture, all calculations are carried out for the International Standard Atmosphere (ISA) conditions, which makes it possible to compare the results of calculations and flight tests of several aircraft conducted in different climatic zones by recalculating the test results to the parameters of the International Standard Atmosphere, “placing” all aircraft in the same conditions - ISA conditions. Historically, the materials for the development of international standards in the field of ISA were developed in the USSR and formed the basis of international standards ISO, which in turn became the basis of documents ICAO 7488/3 of the International Civil Aviation Organisation (ICAO) and subsequently became state documents on standardisation identical to international standards ISO. The above works were carried out within the framework of the international technical subcommittee on standardisation ISO/TC 20/SC 6 “standard atmosphere”, formed in 1980 as part of the international standardisation body ISO/TC 20/SC 6 “standard atmosphere”. Standardisation is the basis for the unity of measurements. In the Russian Federation, the technical committee on standardisation TC 484 “Standard Atmosphere” was established. The developed model of the International Standard Atmosphere allows us to compare the results of calculations and flight tests of several aircrafts conducted in different climatic zones by recalculating the test results to the parameters of the International Standard Atmosphere, “placing” all aircrafts in the same conditions - ISA conditions. With the development of aviation and space technology, the list of atmospheric parameters subject to normalisation and standardisation is expanding.

Effect of flight level to maximum power utilization for PEMFC/supercapacitor hybrid uav with switched reluctance motor thruster

In this study, the performance of a fixed-wing PEMFC/SC hybrid UAV with an SRM thruster is investigated by different flight levels. Flight levels negatively affect the PEMFC due to changing ambient conditions such as air temperature and pressure. Hence, the international standard atmosphere (ISA) model for PEMFC voltage-current-power ratings is derived with 100 m intervals for 0–7500 m. The demanded electrical power by UAV increases at high altitudes due to the compressor used for air pressurization of PEMFC. Therefore, a maximum power point tracking (MPPT) model of PEMFC should be utilized in case of overshooting the maximum power level that can be provided. While the main power source of the UAV is PEMFC, the auxiliary power source is SC. The SC is preferred for DC bus voltage stabilization, avoiding maximum power point overshoot, and preventing high instantaneous current that shortens the lifespan of PEMFC. The proposed UAV model is verified at MATLAB/Simulink and the results prove that the system is satisfying.

Physical Modeling, Simulation and Validation of Small Fixed-Wing UAV

The flight dynamics of small Unmanned Aerial Vehicles (UAVs) exhibits substantial nonlinear features which should not be ignored in simulation or analysis. In this work, a complete six degrees of freedom (6DOF) and high-fidelity simulation model of a case study of a small fixed-wing UAV is developed. To accurately estimate the nonlinear UAV’s mathematical model, the mass-inertia properties are investigated through experiments. Moreover, the aerodynamic model characteristics are estimated using a semi-empirical technique to estimate the aerodynamic coefficients and derivatives. The propulsion system’s physical characteristics are estimated and analyzed through experimental measurements. The International Standard Atmosphere (ISA) tables are used to develop the atmospheric model. Furthermore, the actuation model is estimated and validated experimentally through system identification techniques. These estimated sub-model data are integrated to complete the nonlinear flight dynamics model in MATLAB® (Simulink). The UAV trim calculations are graphically derived and then the developed model is tested and verified. The open loop simulation results have proved the superiority of the utilized model techniques to describe the UAV motion. Moreover, it gives very optimistic results to test the flight dynamics of the UAV and design a consistent flight control and guidance system.

Influence of Aviation Conditions on Lithium-Ion Batteries

Stopping climate change and global temperature rise needs a determined course of action in several technology fields and the development of ZERO-CO2 aircrafts and low emission battery-ICE aircrafts, as well. Batteries in hybrid-electric or fully-electric aircrafts are an option to decrease emissions. The battery requirements for the aviation sector regarding specific energy, energy density and safety are high. Lithium-ion batteries are today the only commercially available technology to meet these requirements, but adequate technology readiness is only proven for the automobile sector while TRL levels 8 and 9 for implementation readiness of batteries in the aviation sector is still ongoing. A crucial aspect for aviation applications is the operation behavior in high altitude including low temperature, wide humidity ranges and especially low pressure. There is little published information available on the effects of low pressure on battery performance. Therefore, different commercially available lithium-ion batteries were tested from atmospheric pressure down to 250 hPa, which corresponds to a flight altitude up to 10,000 m, according to the International Standard Atmosphere (ISA). Since NMC and NCA cathode materials are promising for aviation due to their high specific energy, and LFP cells are promising because of their high safety, the following batteries were selected: 20 Ah HP602030 LFP cylindrical cell (EAS Batteries GmbH, GER). 10 Ah HE 341440 NCA cylindrical cell (EAS Batteries GmbH, GER). 12 Ah SLPB065070180 NMC pouch cell (Kokam Co. Ltd., KOR) To examine the influence of pressure on the battery behavior, constant current discharge tests at 1C 0.5C and 2C rates as well as impedance measurements in 10% SOC steps were performed in a climate controlled low-pressure chamber. The batteries were operated under different temperatures and in controlled low pressure. The pressure was first set to atmospheric pressure as a reference and then lowered to 750 hPa, 500 hPa and 250 hPa. To further examine altitude influences, current cycles and impedance spectroscopy were performed at different temperature and humidity values.The behavior of the cylindrical cells showed hardly any pressure dependency, while the examined pouch cell’s behavior did show a pressure dependency.All batteries showed strong temperature dependency but hardly any humidity dependency.In this presentation the influence of pressure, temperature and humidity on voltage, capacity, resistance and impedance of the afore mentioned batteries will be discussed. The batteries are compared regarding their aviation applicability with special respect to low pressure. The parameters for each environmental condition can then be used to parameterize a battery model suitable for aviation applications.

SimEx: A Tool for the Rapid Evaluation of the Effects of Explosions

The dynamic response of structural elements subjected to blast loading is a problem of growing interest in the field of defense and security. In this work, a novel computational tool for the rapid evaluation of the effects of explosions, hereafter referred to as SimEx, is presented and discussed. The classical correlations for the reference chemical (1 kg of TNT) and nuclear (106 kg of TNT) explosions, both spherical and hemispherical, are used together with the blast wave scaling laws and the International Standard Atmosphere (ISA) to compute the dynamic response of Single-Degree-of-Freedom (SDOF) systems subject to blast loading. The underlying simplifications in the analysis of the structural response follow the directives established by UFC 3-340-02 and the Protective Design Center Technical Reports of the US Army Corps of Engineers. This offers useful estimates with a low computational cost that enable in particular the computation of damage diagrams in the Charge Weight–Standoff distance (CW–S) space for the rapid screening of component (or building) damage levels. SimEx is a computer application based on Matlab and developed by the Fluid Mechanics Research Group at University Carlos III of Madrid (UC3M). It has been successfully used for both teaching and research purposes in the Degree in Security Engineering, taught to the future Guardia Civil officers at the Spanish University Center of the Civil Guard (CUGC). This dual use has allowed the development of the application well beyond its initial objective, testing on one hand the implemented capacities by undergraduate cadets with the end-user profile, and implementing new functionalities and utilities by Masters and PhD students. With this experience, the application has been continuously growing since its initial inception in 2014 both at a visual and a functional level, including new effects in the propagation of the blast waves, such as clearing and confinement, and incorporating new calculation assistants, such as those for the thermochemical analysis of explosive mixtures; crater formation; fragment mass distributions, ejection speeds and ballistic trajectories; and the statistical evaluation of damage to people due to overpressure, body projection, and fragment injuries.

Design of a Battery Cooling System for Hybrid Electric Aircraft

A battery thermal management system (BTMS) for a hybrid electric aircraft is designed. Hot-day takeoff conditions are assumed, resulting in an ambient temperature higher than the allowed battery temperature. Thus, a heat pump has to be employed in the BTMS. All available heat pump technologies are assessed and compared. In a qualitative downselection process, thermoelectric cooling is chosen as heat pump technology. A computational model of the thermoelectric modules (TEMs), as well as a finned ram air heat exchanger with a puller fan (HEX) as heat sink, and a simple heat pipe model for heat acquisition from the battery are developed. Sensitivity analyses of the TEM and HEX models are conducted for comprehension of the relevant design variables. Finally, a BTMS is designed and optimized for a 19-seat hybrid electric aircraft with an all-electric design mission and a combustion engine for range extension. The temperature deviation from the International Standard Atmosphere, the battery operating temperature, and the battery efficiency span a three-dimensional design space rather than being defined to three discrete values. The most adverse combination of these three parameters results in an optimized BTMS that increases the maximum takeoff mass of the aircraft by 16% without the consideration of any snowball effects, whereas the most advantageous combination leads to a 2% increase.

К вопросу об определении барометрической высоты механическим альтиметром и системой воздушных сигналов

Reliable information on the altitude and speed parameters of the aircraft (AC) is required for safety and increasing the efficiency of flight missions. In this regard, the task of studying the features and improving the characteristics of existing meters for these parameters is quite urgent. Purpose of the study. The features of determining the barometric altitude of an aircraft by a mechanical indicator and an air signal system (SHS) are considered. Materials and methods. The essence of the baro¬metric method for measuring the altitude of the aircraft is analyzed, it is found that to build a clear theory of the method, it is necessary to know the dependence of the temperature of the atmosphere on the altitude. Moreover, it is known that this temperature usually decreases with increasing altitude. However, the rate of decrease unpredictably changes both at different times of the year and day, and at different points and at different heights. Moreover, in some cases, in a certain range of heights, the temperature may not decrease, but, on the contrary, increase. In this regard, there is some problem with the formalization of this dependence. When constructing a theory of the method for mechanical pointers, one proceeds in such a way that the dependence for the International Stan¬dard Atmosphere (ISA) is used, in which temperature and pressure are uniquely related. When developing algorithmic support for SHS, it is considered that the temperature decreases with height as well as in a standard atmosphere, but its actual value at the level of the reference plane can be anything that is not at all related to pressure and altitude in ISA. For this purpose, a temperature control unit is provided in the SHS at the level of the height reference plane. Results. A formula has been obtained that makes it possible to determine the difference between the readings of a mechanical altimeter and SHS. It was found that this difference is a function of pressure and temperature at the level of the plane of the altitude reference plane, as well as the actual pressure at the flight altitude. Numerical studies have been carried out to assess the discrepancy between the readings. It has been established that the relative difference in readings can reach the order of magnitude under ope¬rating conditions, and it practically does not depend on the aircraft altitude and is, in essence, a constant for the indicated initial values of pressure and temperature (at the level of the plane of the altitude reference). Conclusion. The developed algorithms can be used to improve the algorithmic support of the SHS.

Evaluation of a small drone performance using fuel cell and battery; Constraint and mission analyzes

Drones have important uses such as relief, photography, military purposes. However, designing an efficient propulsion system for the drone that can ensure its reliability and flight endurance is a major challenge. Relying on an all-electric propulsion system can enhance flight endurance and improve drone performance. In this article, two types of fuel cells, proton-exchange membrane fuel cell and solid oxide fuel cell have been considered (individually) to supply the power consumption of the drone. In addition, a battery is embedded in parallel with the fuel cell in the propulsion system. To evaluate the performance of the drone, two constraint and mission analyzes have been established to determine the power consumption of the drone during various flight modes and weight distribution of the drone. Then a comparison on the drone’s performance with respect to the electrochemical behavior of both fuel cells is provided. Finally, the performance of fuel cells in propulsion system based on the International Standard Atmosphere model is discussed. Results revealed that the solid oxide fuel cell and proton-exchange membrane fuel cell can generate 82.1 and 162.7 W of electrical power respectively. In addition, drone needs about 1.46 kW of electrical power. The flight endurance and total weight of the drone are equal to 6.8 h and 6.3 kg, respectively. It was also found that the required area of the solid oxide fuel cell is approximately 98% less than the proton-exchange membrane fuel cell to provide the total electricity consumption of the drone. Furthermore, the effect of increasing cruise speed on drone’s net weight is more remarkable compared to power consumption.

Use of the Speed of Sound for Heights Below the Troposphere in Colombia to Determine the Distance of Lightning Strikes

Abstract Objective: This paper aims to conduct a study of the speed of sound in Colombia for heights below the troposphere, taking into account the temperature as a function of altitude for a tropical zone, in order to determine the distance of lightning strike. Methodology: A speed of sound profile for Colombia was determined by analyzing data on altitude, temperature, and relative humidity in the country. The temperature equation as a function of altitude presented by NASA and the International Standard Atmosphere was taken into account to compare with the data obtained from meteorological stations located in different parts of the country (this stations measure information of temperature and other physical variables). Additionally, a sound profile map of Colombia was obtained using the interpolation in ArcGIS software. Results: A map of the speed of sound in Colombia was obtained, thus finding the speed range in the country. Likewise, the temperature was determined as function of the altitude for Colombia (a country in the tropical zone), with which it is possible to have data closer to the real speed of sound in the place and determinate the distance of impact of a lighting in land. Conclusions: The study of variables such as temperature, relative humidity, and altitude helped determine the speed of sound for Colombia. It was also possible to define a temperature equation as a function of altitude for a tropical country; with the speed of sound, the radius of impact of a lightning can be found.

An estimation method for the fuel burn and other performance characteristics of civil transport aircraft in the cruise. Part 1 fundamental quantities and governing relations for a general atmosphere

AbstractThis paper is one of a series addressing the need for simple, yet accurate, methods for the estimation of cruise fuel burn and other important aircraft performance parameters. Here, a previously published, constant Reynolds number model for turbofan-powered, civil transport aircraft is extended to include Reynolds number effects. Provided the variation of temperature with pressure is known, the method is applicable to flight in any atmospheric conditions. For a given aircraft, cruising in a given atmosphere, there is a single Mach number and Flight Level pair, at which the fuel burn per unit distance travelled through the air has an absolute minimum value. Both these quantities depend upon the Reynolds number, which, in turn, depends upon the aircraft weight and the atmospheric vertical temperature profile. Simple, explicit expressions are developed for all parameters at the optimum condition. These are shown to be in close agreement with numerical solutions of the governing equations. It is found that typical operational mass and temperature profile variations can change cruise fuel burn rate by several percent. In the International Standard Atmosphere, when the speed and altitude deviate from their optimum values, the fuel burn penalty is reduced slightly relative to the constant Reynolds number case. By way of example, the method is used to estimate the minimum fuel, speed-versus-height trajectory for cruise in a realistic atmosphere.For each aircraft, cruise fuel burn is found to be governed by six independent parameters. All are constants. Two are simple, involving only size and weight, whereas four are complex and must be determined by either theoretical, or empirical, means. The estimation of these quantities will be considered in Part 2.

Aircraft Emission Inventory and Characteristics of the Airport Cluster in the Guangdong–Hong Kong–Macao Greater Bay Area, China

In this study, a compound method using modified Boeing Fuel Flow Method 2 (BFFM2) and an updated First Order Approximation V3.0 (FOA3.0) method deploying the ICAO Time-in-Mode (TIM) was used to produce a more reliable aircraft emission inventory for the Guangdong–Hong Kong–Macao Greater Bay Area (GBA). The results show that compared with the International Standard Atmosphere (ISA) conditions, the total emission of nitrogen oxides(NOx) decreased by 17.7%, while carbon monoxide(CO) and hydrocarbons(HC) emissions increased by 11.2%. We confirmed that taxiing is the phase in which an aircraft emits the most pollutants. These pollutant emissions will decrease by 0.3% to 3.9% if the taxiing time is reduced by 1 minute. Furthermore, the impact of reducing taxi-out time on emissions is more significant than that of reducing the taxi-in time. Taking the total aircraft emission factors as the main performance indicators, Hong Kong International Airport (VHHH) contributes the most to the total emissions of the GBA, while the Zhuhai airport(ZGSD) contributes the least. The contribution of an individual airport to the total emissions of the GBA is mainly determined by the proportion of Boeing B77L, B77W, and B744.

К настройке алгоритма вычисления барометрической высоты летательного аппарата системой воздушных сигналов

Introduction. In conditions of increasing requirements for the accuracy of determining the altitude and speed parameters of aircraft, the task of tuning up the algorithmic support of their air signal systems (SHS) based on real flight data is quite relevant. Aim. Consider the task of assessing the real value of the temperature gradient along the height of the atmosphere for the purpose of operational adjustment of the standard algorithmic support of SHS. Materials and methods. The traditional approach is to use the numerical value of the temperature gradient, adopted for the international standard atmosphere (ISA). A feature of the proposed approach is that the temperature gradient in height is determined by current measurements of atmospheric pressure and temperature, and not height and temperature. Results. Adaptive and non-adaptive parametric identification algorithms are developed. An improved version of the adaptive algorithm is proposed, which ensures proper tuning of its filtering properties to a specific statistical situation. Algorithms were tested both according to the characteristics of the ISA using computer-generated random interference, and according to the results of real measurements known in the literature obtained using a meteorological rocket. The properties of adaptive and non-adaptive algorithms are compared. It is shown that repeated “running” of the same experimental data through an adaptive algorithm leads to almost the same accuracy of the gradient estimate, as when using a non-adaptive algorithm. Conclusion. The implementation of the algorithms improves the accuracy of determining the barometric altitude of the aircraft (AC), does not require any modernization of the hardware of the SHS, only some addition of the algorithmic support of the system is necessary. The adaptive algorithm is recommended for use in the case of a significant drift of atmospheric characteristics for the operational adjustment of the system.

Performance prediction of aircraft gasoline turbocharged engine at high-altitudes

One-dimensional modeling is considered as one of the most approved methods of studying the characteristics and performance of aircraft reciprocating engines, operating under different flight conditions. Owing to the recent progresses, the one-dimensional modeling of piston engine components including air filter, carburetor, air manifold, inlet and outlet valves, cylinder, piston, crankshaft, exhaust manifold, and turbocharger have become possible using some of the engineering software such as the GT-Power. The present paper is devoted to extensive performance analysis of one of the most widely used aircraft engines, the Rotax-914 which is a spark-ignition (SI) turbocharged aircraft engine, during the climb phase. The survey is conducted through the preparation of a one-dimensional model of the Rotax-914 using GT-power software. Then, to be more consistence with the real flight condition, its turbocharger is accurately modeled by considering the performance curves provided by the manufacturer. Next, to ensure its accuracy, the developed model was validated against the manufacturer data. The important engine performance parameters and turbocharger performance curves were analyzed for different flight aspects up to 30,000 feet. The obtained results indicated that the turbocharger prevents a dramatic decline in engine performance parameters up to 18,000 feet. Thereafter, the compressor reaches the choking region, fails to supply air demand, leading to extreme performance loss. The effect of ambient temperature difference from the international-standard-atmosphere (ISA) on engine performance was also investigated. A 40 °C increase in the ambient temperature causes a 10% drop in the brake horsepower and consequently, a performance loss was observed during the climb. Ambient temperature and altitude differently affect engine performance. The effect of ambient temperature on the engine performance was greater as the compressor reaches the choking region. Beyond the choking region, altitude variations had a greater impact on engine performance than the temperature variations.

Optimization of the flight program of a passenger aircraft taking into account operational limitations and the influence of climatic factors.

The paper presents the results of optimization of the flight program for a passenger aircraft. In mathematical modeling, real operational limitations and possible changes in atmospheric parameters were taken into account. The mathematical model is characterized by the description of the flight of an airplane as a material point and the description of a gas turbine engine by a model of the first level of accuracy. The following operational limitations were considered: the possibility of cruising at certain levels, the restriction on the vertical speed with a decrease, the possibility of transition to the next echelon with an available stock of thrust of 20%. The optimization of the flight program is based on the criterion of the minimum amount of fuel expended for the flight, at a given range. Calculations were conducted for several standards of air temperature change, (depending on the climatic zone). A comparative analysis of the obtained optimization results is carried out, the degree of influence of the change in atmospheric conditions is estimated. When comparing flights in different climatic zones, the fuel costs were compared for a flight program optimized for the given atmospheric conditions and for a flight program optimized for use in the mathematical model of the International Standard Atmosphere (ISA).

Computational Fluid Dynamics Investigation of a Core-Mounted Target-Type Thrust Reverser—Part 2: Reverser Deployed Configuration

Core-mounted target-type thrust reverser (CMTTTR) design was proposed by NASA in the second half of the 90 s. NASA carried out several experiments at static conditions, and their acquired results suggested that the performance characteristics of the CMTTTR design fall short to comply with the mandatory thrust reverser (TR) performance criteria, and were therefore regarded as an infeasible design. However, the authors of this paper believe that the results presented by NASA for the CMTTTR design require further exploration to facilitate the complete understanding of its true performance potential. This part 2 paper is a continuation from Part 1 (reverser stowed configuration) and presents a comprehensive three-dimensional (3D) computational fluid dynamics (CFD) analyses of the CMTTTR in deployed configuration. The acquired results are extensively analyzed for aforementioned TR configuration operating under the static operating conditions at sea level, i.e., sea-level static, International Standard Atmosphere (ISA); the analyses at forward flight conditions will be covered in part 3. The key objectives of this paper are: First, to validate the acquired CFD results with the experimental data provided by NASA; this is achieved by measuring the static pressure values on various surfaces of the deployed CMTTTR model. The second objective is to estimate the performance characteristics of the CMTTTR design and corroborate the results with experimental data. The third objective is to estimate the pressure thrust (i.e., axial thrust generated due to the pressure difference across various reverser surfaces) and discuss its significance for formulating the performance of any TR design. The fourth objective is to investigate the influence of kicker plate installation on overall TR performance. The fifth and final objective is to examine and discuss the overall flow physics associated with the thrust reverse under deployed configuration.