Reference Temperature Method Research Articles

Thermo-aeroelastic analysis of triangular composite plate with embedded shape memory alloy at supersonic flow

This work investigates the effectiveness of a shape-memory alloy (SMA) in controlling the instabilities of triangular composite plates under supersonic flow. Lagoudas’ quadratic polynomial hardening theory models the SMA effect. First-order piston theory was used for the aerodynamic modeling, and the reference-temperature method was used for modeling the thermal heating. The buckling and post-buckling behaviors were studied for different boundary conditions with four different layups. In addition, buckling and post-buckling of the composite plate, with and without shape memory alloy wire, has been studied. The effect of SMA wire on aeroelastic instabilities is accurately studied. The embedded SMA wire significantly increased the stability region (postpone divergence and flutter velocities) and buckling temperature. Also, the time responses of the triangular composite plate are determined at different Mach numbers, showing that by increasing the Mach number, the SMA wire can control or decrease the vibration amplitudes.

Transient Aerodynamic Heating Effects on Aeroelastic Dynamic Response of Curved Fiber-Reinforced Composite Panel in Supersonic Airflow

The nonlinear aero-thermo-elastic behaviors of the curved-fiber composite panel are studied accounting for transient aerodynamic heating effects with two-way coupling method. The Von Karman assumption is utilized to depict the large deflection of the composite panel and the supersonic aerodynamic load is incorporated through first-order piston theory, respectively. The aerothermal and aeroelastic models are coupled with each other by performing the iterative procedure in the time domain. The aerodynamic heating heat flux is obtained by the Eckert reference temperature method, the transient temperature field is solved by the finite difference method, and the aeroelastic dynamic response is solved by the combination of the Newmark method and Newton–Raphson method. The accuracy and validity of the established model are justified by comparing the present solutions with the present numerical results in published literature. Then, the influences of the key parameters such as the transient temperature field, coupling step, heat transfer coefficient, shock wave, initial disturbance force and fiber configurations on aerothermoelastic responses of the composite panel respectively are discussed in detail.

Thermal‐Infrared Emission Spectroscopy of Graybody Minerals (Sulfide): Implications for Extraterrestrial Exploration for Magmatic Ore Deposits

AbstractGraybody materials exhibit systematically low emissivity across their spectrum. This characteristic violates the key assumption of unit emissivity at some wavenumber in the spectrum used to calibrate thermal‐infrared emission data. This assumption makes graybody materials “appear colder” than their actual physical temperature and imparts a slope in emission spectra that is non‐physical in nature, both of which affect interpretations of planetary surfaces. Pyrrhotite derived from the Stillwater Complex's J‐M Reef in Montana, USA exhibits systemic graybody behavior across its mid‐infrared spectrum and thus has a steep negative spectral slope from high to low wavenumbers when calibrated using conventional methods. A new measurement technique is introduced for deriving the absolute emissivity of graybody materials using reference samples with known Christiansen Frequencies during calibration. The reference temperature method significantly reduces the spectral slope of and provides a more accurate estimation of the absolute emissivity of graybody materials. After correcting the temperature of pyrrhotite using results from a series of reference experiments, we conclude that the emission spectrum of pyrrhotite is spectrally featureless and has a maximum emissivity of ∼0.7. If sulfide mineral deposits are exposed on Mars, they will not be identified using spectral features found in the mid‐infrared (5–40 μm). However, they could be located by identifying basaltic terrain that appears colder than their surroundings and with apparent emissivity spectra that exhibit negative spectral slopes from high to low wavenumbers and are spectrally neutral.

Fluid-thermo-structural response of actively cooled scramjet combustor in hypersonic accelerating-cruise flight

This paper presents a three-dimensional transient fluid-thermo-structural study of an actively cooled sandwich panel under hypersonic accelerating-cruise flight conditions. The thermo-structural loads are estimated using a high-speed gas-dynamic flow model combined with Eckert's reference temperature method. The thermal safeguard capacity with endothermic fuel-based active cooling shows design optimization scope compared to the passive system. A passive system has to survive excessively high temperatures with a less severe thermal gradient over the panel thickness and a high heat leakage into the structural interior or back wall. The actively cooled system with fuel develops significant panel bending due to a through-thickness thermal gradient in the cooling channels. For a Mach 7 flight, an actively cooled system reduces the bending deformation by 90% and improves the thermal safeguard capacity by 70% compared to a passive approach. Active cooling can effectively control the excessive thermo-structural deformation and fuel heating below its cracking temperature to improve combustion. Our parametric design study indicates the influence of fuel as a coolant for increasing combustion efficiency and the associated complexity and trade-off in the heat transfer and thermo-structural deformation behavior. This transient response under hypersonic accelerating-cruise flight conditions requires a careful review of the scramjet engine thermo-structural design philosophy. The design approach has applications in bodies encountering highly transient thermo-structural loads with differently purposed fuels/coolants.

Field test of two background temperature correction methods of dual probe heat pulse method

Dual probe heat pulse (DPHP) method is a useful method to measure many soil physical properties in-situ, which play important roles in soil science, meteorology, hydrology and so on. It is essential to remove the effect of background temperature fluctuations for DPHP estimations. We compared two background temperature correction methods of DPHP method, linear trend method (LTM) and reference temperature method (RTM) in field experiments. We found that misused LTM lead to large errors in DPHP estimations with shallow burial depth. When the burial depth of temperature probe (d) of DPHP sensor is <14 mm, background temperature changes are related to environmental factors such as solar radiation and wind, and RTM can decrease the relative error (RE) in thermal conductivity (λ) from 667% to 68%. Besides, our results also showed during the period from 9:00 to 16:00, the RTM method is valuable for getting accurate DPHP estimations with the 2 mm burial depth. After replacing LTM by RTM, REs were decreased to (6 ± 28)% from (17 ± 87) % in λ, and were decreased to (9 ± 17)% from (10 ± 18)% in volumetric heat capacity (Cv). But RTM is also affected by other factors such as the spatially heterogeneous temperature distribution of tested sample and should be taken into account in the further studies.

Control-oriented quasi-one dimensional modeling method for scramjet

Abstract The scramjet model used in the control system needs to consider both accuracy and real-time requirements. In this paper, a quasi-one dimensional component level model method for scramjet is proposed. Under the premise of meeting the real-time requirement, this modeling method has higher calculation accuracy, and can obtain higher thrust performance without exceeding the limits. In order to improve the calculation accuracy of the scramjet model, the boundary layer correction and the interaction between shock wave and boundary layer are considered in the inlet and isolator modeling process by using reference temperature method. For the scramjet compression components, the inlet and isolator models are simplified based on shock train theory to improve real-time performance. Meanwhile, the combustor model and nozzle model are established and the combustor model is established with the combination of equal area and isothermal based on temperature limitation. Considering the boundary layer correction, the accuracy of the inlet and isolator model is improved by 8.79 and 7.9% respectively. Compared with the equal area combustion mode, the combined combustion mode can get higher thrust. In the C + + simulation environment, the average calculation time of each state point is about 25 ms which meets the real-time requirements.

Research on Calculation Method of External Convective Heat Transfer Coefficient of Civil Aircraft

Based on the study of calculation method of external convective heat transfer coefficient of civil aircraft, this paper proposes a simulation calculation method, heat transfer formula method, which can be used to solve the external convective heat transfer coefficient of each part of aircraft under flight conditions. The STAR-CCM+ software is used as the calculation platform. Based on a civil aircraft, the theoretical reference temperature method is used to verify the heat transfer formula method. The results show that the difference between the simulation value of heat transfer formula method and the theoretical value of reference temperature method is about 8% of the theoretical value, so the value calculated by heat transfer formula method is basically accurate. Correct the theoretical value based on the simulated value, that can be used to solve the external convective heat transfer coefficient of each part of aircraft.

Efficient Prediction of the Temperature History of a Hypersonic Vehicle Throughout the Mission Trajectory with an Aerodynamic Thermal Load Element

This work introduces an efficient approach for aero-thermo-mechanical analysis to predict the temperature history of a hypersonic vehicle during its full mission trajectory. The approach uses a recently proposed aerodynamic thermal load element, in which the effects of aerodynamic pressure and aerodynamic heating are efficiently considered using local piston theory and the Eckert reference temperature method, respectively. This element is implemented as a user-subroutine in commercial software to handle realistic models for the aero-thermo-mechanical analysis, such as the X-43A. A finite element model of the X-43A is constructed for a benchmark test. Using the model, an aero-thermo-mechanical analysis is carried out while considering the full mission trajectory of the X-43A. The predicted temperature results are compared with recorded flight test data from the X-43A. Additionally, the aero-thermo-mechanical behavior of the hypersonic vehicle is investigated according to various parameters. The investigation confirms that the approach can be efficiently used in the design of hypersonic vehicles with considerable savings in computational cost.

Performance improvement in double-ended RDTS by suppressing the local external physics perturbation and intermodal dispersion

We propose and experimentally demonstrate a novel Raman-based distributed fiber-optics temperature sensor (RDTS) for improving the temperature measurement accuracy and engineering applicability. The proposed method is based on double-ended demodulation with a reference temperature and dynamic dispersion difference compensation method, which can suppress the effect of local external physics perturbation and intermodal dispersion on temperature demodulation results. Moreover, the system can omit the pre-calibration process by using the reference temperature before the temperature measurement. The experimental results of dispersion compensation indicate that the temperature accuracy optimizes from 5.6°C to 1.2°C, and the temperature uncertainty decreases from 16.8°C to 2.4°C. Moreover, the double-ended configuration can automatically compensate the local external physics perturbation of the sensing fiber, which exhibits a distinctive improvement.

Control-relevant modeling and performance limitation analysis for flexible air-breathing hypersonic vehicles

Overall design in many air-breathing hypersonic vehicles (AHSVs) uses the integrated airframe/scramjet configuration formed with elongated fuselage and light structure, resulting in intricate aerodynamic–propulsive-elastic couplings. In this paper, a high-fidelity flexible AHSV dynamic modeling procedure is proposed at first, which consists of three main parts: i) panel method based aerodynamic modeling; ii) scramjet modeling; iii) finite-element-method (FEM) based aeroelastic modeling. Unsteady and viscous effects are estimated through linear piston theory and Eckert's reference temperature method, respectively. Additional models for shock interaction, ram-to-scram transition, and finite rate chemistry reaction are also included. To overcome the intractability of physics-based model, the Morris method is applied to identify the main influence factors during surrogate modeling. Then to investigate the flexible mode effect on the stability boundary, the principle of Null-controllability region (NCR) is employed, under control input saturation and bandwidth constraints.The X-43A like air-breathing hypersonic vehicle operating at cruise condition is studied for verification purpose, and the Monte-Carlo simulation results demonstrate the effectiveness of the proposed dynamic modeling and performance limitation analysis approach.

Studies for aeroelastic characteristics and nonlinear response of FG-CNT reinforced composite panel considering the transient heat conduction

A solution of computing the aero-thermal-elastic problem of FG-CNT reinforced composite panel is proposed, in which the aerodynamic heating and transient heat conduction are accounted for. The Von-Karman theory and the third-order piston theory are employed to establish the motion equation of FG-CNT panel, where the thermo-mechanical properties of polymer matrix and carbon nanotubes are set to be location- and temperature-dependent. The Ecker’s reference temperature method and heat flux equation are adopted to calculate the heat flux. Two dimensional transient heat conduction including in-plane and through-thickness directions are computed using a finite difference approach. In the calculation process, the influence of elastic deformation on the aero-heating and aero-heating on panel stiffness is included. The aeroelastic stability analysis is presented for the influential parameters such as CNT volume fraction, distribution and aspect ratio of the panel. Moreover, the effects of the oblique shock wave parameters on the critical flutter pressure and dynamic response are discussed. Numerical results show that the coupled solution leads to the inhomogeneous aero-heating, which results in nonuniform temperature distributions and material property degradations. Consequently, the stability boundary and dynamic response are varied with respect to the case without considering the coupling effect.

Design of a three-dimensional scramjet nozzle considering lateral expansion and geometric constraints

A new method based on quasi two-dimensional supersonic flow and maximum thrust theory to design a three-dimensional nozzle while considering lateral expansion and geometric constraints is presented in this paper. To generate the configuration of the three-dimensional nozzle, the inviscid flowfield is calculated through the method of characteristics, and the reference temperature method is applied to correct the boundary layer thickness. The computational fluid dynamics approach is used to obtain the aerodynamic performance of the nozzle. Results show that the initial arc radius slightly influences the axial thrust coefficient, whereas the variations in the lateral expansion contour, the length and initial expansion angle of the lower cowl significantly affect the axial thrust coefficient. The three-dimensional nozzle designed by streamline tracing technique is also investigated for comparison to verify the superiority of the new method. The proposed nozzle shows increases in the axial thrust coefficient, lift, and pitching moment of 6.86%, 203.15%, and 642.86%, respectively, at the design point, compared with the nozzle designed by streamline tracing approach. In addition, the lateral expansion accounts for 22.46% of the entire axial thrust, while it has no contribution to the lift and pitching moment in the proposed nozzle.

Design and analysis on three-dimensional scramjet nozzles with shape transition

A trimming method to optimize the configuration of a three-dimensional asymmetric nozzle with shape transition, which aims to obtain good aerodynamic performance and to save weight at the same time, is presented in this paper. Then the effects of the entry shape on the performance of the three-dimensional nozzle are investigated. The streamline tracing involved in an axisymmetric flowfield with optimal thrust is employed to obtain the inviscid contour of the three-dimensional nozzle with shape transition, and the reference temperature method is applied to correct the thickness of the boundary layer. The performance of the designed nozzle is obtained by using computational fluid dynamics. The calculated results show that the trimmed nozzle gains increases in the lift and pitching moment by 427.00% and 10.80%, respectively, with only a 0.76% decrease in the axial thrust coefficient, while the weight can be reduced by as much as 37.51%. For the nozzles with elliptical entrances, as the axial ratio ranges from 1.0 to 2.0, the axial thrust coefficient is increased by 5.38%, while the lift is decreased by 67.74%. When considering the nozzles with rectangular entrances, as the aspect ratio ranges 1.0 to 2.0, the axial thrust coefficient is increased by 3.58%, while the lift and pitching moment are decreased by 82.09% and 16.43%, respectively. Most of the axial thrust is produced on the upper-wall and side-walls in all nozzles, and the contribution of the expansion flow along the side-walls on the thrust generation is pronounced in the nozzle with a relatively smaller entry ratio. However, the majority of the lift and pitching moment are generated on the upper-wall and lower-wall. Besides, the viscosity loss and weight can be reduced by applying the elliptical cross-section in the propulsion system.

A new design method of single expansion ramp nozzles under geometric constraints for scramjets

A new method based on maximum thrust theory to design a two-dimensional single expansion ramp nozzle with geometric constraints directly is presented in this paper. To generate the contour of the nozzle, the inviscid flowfield is calculated by using the method of characteristics and the reference temperature method is applied to correct the boundary layer thickness. The computational fluid dynamics approach is employed to obtain the aerodynamic performance of the nozzles. The results show that the initial arc radius slightly influences the axial thrust coefficient and that the variations in the length and initial expansion angle of the cowl significantly affect the axial thrust coefficient. The nozzle designed by truncating ideal nozzle is also investigated for comparison to verify the superiority of this new method. The nozzle designed by this proposed method shows increases in the axial thrust coefficient, lift, and pitching moment of 5.5%, 1098.2% and 20.3%, respectively, at the design point. Moreover, the proposed nozzle provided the positive lift with considerable increments in the axial thrust coefficient and in the pitching moment at off-design operations.

A high-efficiency aerothermoelastic analysis method

In this paper, a high-efficiency aerothermoelastic analysis method based on unified hypersonic lifting surface theory is established. The method adopts a two-way coupling form that couples the structure, aerodynamic force, and aerodynamic thermo and heat conduction. The aerodynamic force is first calculated based on unified hypersonic lifting surface theory, and then the Eckert reference temperature method is used to solve the temperature field, where the transient heat conduction is solved using Fourier’s law, and the modal method is used for the aeroelastic correction. Finally, flutter is analyzed based on the p-k method. The aerothermoelastic behavior of a typical hypersonic low-aspect ratio wing is then analyzed, and the results indicate the following: (1) the combined effects of the aerodynamic load and thermal load both deform the wing, which would increase if the flexibility, size, and flight time of the hypersonic aircraft increase; (2) the effect of heat accumulation should be noted, and therefore, the trajectory parameters should be considered in the design of hypersonic flight vehicles to avoid hazardous conditions, such as flutter.

Effect of Variable Properties Within a Boundary Layer With Large Freestream-to-Wall Temperature Differences

Modern gas-turbine engines are characterized by high core-flow temperatures and significantly lower turbine-surface temperatures. This can lead to large property variations within the boundary layers on the turbine surfaces. However, cooling of turbines is generally studied near room temperature, where property variation within the boundary layer is negligible. The present study first employs computational fluid dynamics to validate two methods for quantifying the effect of variable properties in a boundary layer: the reference temperature method and the temperature ratio method. The computational results are then used to expand the generality of the temperature ratio method by proposing a slight modification. Next, these methods are used to quantify the effect of variable properties within a boundary layer on measurement techniques, which assume constant properties. Both low-temperature flows near ambient and high-temperature flows with a freestream temperature of 1600 K are considered under both laminar and turbulent conditions. The results show that variable properties have little effect on laminar flows at any temperature or turbulent flows at low temperatures such that constant property methods can be validly employed. However, variable properties are seen to have a profound effect on turbulent flows at high temperatures. For the high-temperature turbulent flow considered, the constant property methods are found to overpredict the convective heat transfer coefficient by up to 54.7% and underpredict the adiabatic wall temperature by up to 209 K. Utilizing the variable property techniques, a new method for measuring the adiabatic wall temperature and variable property heat-transfer coefficient is proposed for variable property flows.

Simulation Analysis of Viscous Effects on Dynamic Performance of Hypersonic Vehicle

In this paper, viscous effects on hypersonic vehicle are investigated. Firstly, viscous aerodynamic is modeled based on reference temperature method, and viscous aerodynamic lift, drag and pitching moment are calculated by the model. And then, aerodynamic forces and elastic deformation are analyzed by comparing with viscous effects on/off. In final, considering the thermal choking condition, longitudinal velocity and attitude boundary values are discussed.

Aerothermoelastic analysis of nonlinear composite laminated panel with aerodynamic heating in hypersonic flow

Aero-thermal-elastic characteristics of hypersonic nonlinear laminated panels considering the shock wave (SW) as well as the aerodynamic heating are studied in this paper. The linear and third order piston theories are used to evaluate the aerodynamic pressure. Hamilton’s principle is employed to establish the aeroelastic model of the laminated panel. In the aerothermal model, the heat flux in the aerodynamic heating is computed using the reference temperature method, and the finite element method (FEM) is applied to solve the transient heat transfer in the panel. After solving the aeroelastic and aerothermal models separately, the solutions are substituted into each other. Then performing the procedure of iteration, the coupling effects among the structure, fluid and thermal are realized. Time domain responses of the laminated panel are obtained. The influences of the shock wave on the aeroleastic characteristics of the panel are also investigated. From the numerical simulation, it can be observed that the aeroelastic stability is weakened with the increase of the temperature change. It is also noted that in the calculation of the airflow parameters for the aerothermoelastic analysis in the hypersonic flow, the influences of the shock wave should be taken into account.

Heat-Flux Measurements with Temperature-Sensitive Paint in a Mach-6 Quiet Tunnel

A procedure based on an analytical inverse method is developed to calculate heat-flux distributions on a model using a time sequence of temperature-sensitive-paint images acquired in the Purdue quiet Mach-6 Ludwieg tube. Temperature-sensitive-paint measurements are conducted on a 7-deg half-angle sharp circular metal cone at Mach 6 to evaluate the accuracy of this method. It is found that the historical effect of the warming-up process of a model in sequential runs becomes significant. Therefore, it must be corrected by in situ calibration to obtain accurate heat-flux distributions. After this historical effect and the effect of the nonuniform coating thickness and other potential factors are corrected, the temperature-sensitive-paint-derived heat-flux distributions on the cone are in good agreement with those given by the similarity solution and the reference temperature method. The error sources in temperature-sensitive-paint heat-flux measurements in the Ludwieg tube are discussed.

Aerothermal-aeroelastic two-way coupling method for hypersonic curved panel flutter

A method to compute aerothermal-aeroelastic two-way coupling for hypersonic curved panel flutter is proposed. The aero-thermo-elastic governing equations of a simply-supported two dimensional curved panel are developed based on the von Karman geometrically non-linear theory. The Galerkin approach is used to simplify the equations into discrete forms, which are solved by the fourth-order Ronger-Kutta method. The third-order piston theory is applied to the aerodynamics. The Eckert’s reference temperature method and the panel heat flux formula are used to compute the aerodynamic heat flux. Several important effects are included, namely 1) two-way coupling considering the effect of elastic deformation on aerodynamic heating and aerodynamic heating on stiffness of structure, 2) accumulation of the aerodynamic heating in real cruise, 3) arbitrary, non-uniform, in-plane and through-thickness temperature distributions, and 4) the effect of initial deformation of curved panel on the flight time to the onset of flutter. Compared with the results of aerothermal-aeroelastic one-way coupling, it is revealed that the two-way coupling which induces decrease of the flight time to the onset of flutter is more dangerous. In addition, importance should be attached to this method in actual analysis.