Supersonic Combustion Flowfield Research Articles

Advances in Femtosecond Coherent Anti-Stokes Raman Scattering for Thermometry

The combustion process is complex and harsh, and the supersonic combustion flow field is also characterized by short duration and supersonic speed, which makes the real-time diagnostic technology for the transient environment extremely demanding. It is of great significance to realize high time-resolved accurate measurement of temperature, component concentration, and other parametric information of the combustion field to study the transient chemical reaction dynamics of the combustion field. Femtosecond CARS spectroscopy can effectively avoid the collision effect between particles in the measurement process and reduce the influence of the non-resonant background to improve the measurement accuracy and realize the time-resolved measurement on a millisecond scale. This paper introduces the development history of femtosecond CARS spectroscopy, points out its advantages and disadvantages, and looks forward to the future development trend to carry out high time-resolved measurements, establish a database of temperature changes in various complex combustion fields, and provide support for the study of engine mechanisms.

Supersonic combustion field evolution prediction in scramjet engine using a deblurring multi-scale attention network

Traditional deep learning technologies often overlook crucial spatiotemporal information when reconstructing supersonic combustion flow fields. These models can only reconstruct the current flow field, resulting in poor visual quality due to the atomization effect. This research introduces experimental studies on scramjet combustion flows with variable injection pressures at the inlet of a scramjet Mach 2.5, creating image datasets of flame evolution processes over different predicted time spans (5, 10, and 15 ms). It develops a multi-scale attention algorithm with deblurring (DB-MSAN) capabilities to predict the flame image after a certain period using the wall pressure signal from the previous moment. The algorithm employs two multi-head attention channels at different scales to extract timing information, encouraging the model to focus on the anisotropy of the input data and to identify more significant features. The DB module utilizes the adapted atmospheric scattering model to jointly estimate the transmission map and atmospheric light, achieving an exceptional defogging effect with a minimal computational expense. The model’s overall performance is assessed regarding efficiency and accuracy, comparing DB-MSAN’s performance differences with those of Chen’s and the ResNet16 models across multi-span, large-span, various test sets, and sparse wall measurement points. Experimental results indicated that, compared to other methods, DB-MSAN achieves a maximum peak signal-to-noise index improvement of 76.10 % and an SSIM index improvement of 155.26 % when predicting average-quality images. The prediction accuracy remains stable, even with sparse pressure input. In addition, DB-MSAN strikes an optimal balance between the number of parameters and model accuracy; it represents only 16.97 % of the volume compared to Chen’s model, which is 585 MB.

An improved flamelet/progress variable modeling in a hydrogen-fueled scramjet

The compressible nature of the supersonic combustion flow field results in significant temperature and pressure fluctuations. In the standard Flamelet/Progress Variable (FPV) model, not only is it difficult to specify the tabulation boundary conditions for the flamelet database, but using a single fixed flamelet database to describe the entire complex flow field will likely introduce significant errors. To account for this, an improved FPV model is proposed in this study. Dynamic adjustment of the tabulation pressure in the flamelet database is achieved by scaling the benchmark flamelet database to match the local pressure. Real-time calibration of the tabulation oxidant temperature and fuel temperature is also achieved by real-time statistics and interpolation between four flamelet databases. This allows the improved FPV model to adaptively match the target supersonic combustion flow field without the difficulty of specifying tabulation pressures and temperatures as in the standard FPV model, and it avoids the limitations imposed by the single flamelet database. An additional species transport equation is also incorporated into this model to improve species prediction. Based on the simulations of a hydrogen-fueled scramjet with low-and-high total temperature, the improved model is shown to better restore wall-pressure profiles, combustion mode predictions and OH (hydroxyl radical) distributions, compared to the standard FPV model and comparable studies. This indicates that the effects of compressibility on the chemical reaction rates have been considered.

Research on time series prediction of the flow field in supersonic combustor based on deep learning

The detection of supersonic combustion flow field provides valuable reference information for the efficient combustion of hypersonic vehicles. From the perspectives of scramjet operation during flight as well as experiment where the measured information is limited, flow field prediction using deep learning technology is a promising method that can provide future flow field evolution in scramjet combustors. This study proposes a convolutional neural network model of multi-temporal pressure information fusion to predict combustion flow field based on wall pressure. In addition, it analyzes the effect of various wall pressures on the flow field prediction. An experimental dataset of supersonic ground-pulse wind tunnel tests with different injection pressures is built through the combustion chamber wall under self-ignition conditions of a hydrogen-fueled scramjet. The results showed that the predicted flow field of the combustor agrees with the ground test, average correlation coefficient reached more than 90% and when the wall pressure number is 32, the average peak signal to noise ratio reaches about 23. and the main wave structure and turbulent boundary layer region are restored even if there is a shock boundary layer interaction.

Evaluation of flamelet/progress variable model for the applications in supersonic combustion using hybrid RANS/LES approach

The Flamelet/Progress Variables (FPV) model coupled with large eddy simulation (LES) has been performed to appraise its applicability and performance for supersonic hydrogen combustion, based on the simulations of strut based scramjet and cavity based combustors. For the two typical schemes involved in the present work, the validation of the numerical solver compiled by in-house code is realized by the comparisons with the experimental measurements including temperature, velocity, and wall pressure. The methods for calculating mixture fraction variance in the framework of LES are also investigated and examined in both cases. It can be observed that the diffusion combustion mode is in the dominant state for the whole reactive zone in both cases, where the FPV model is physically compatible. Through the comparisons of the scale similarity method and the transport equation method for calculating mixture fraction variance, the importance of calculating this sub-grid turbulent scalar flux is emphasized. The results show that the mixture fraction variance calculated by the scale similarity model is locally high, but its diffusion is insufficient, which makes the local temperature of the combustion flow field overpredicted and concentrated. On the contrary, the transport equation of the mixture fraction variance can overcome this problem and improve the combustion prediction accuracy, thus, it is more recommended if the computing resources are guaranteed. Qualitatively, the Flamelet/Progress Variables (FPV) model with the large eddy simulation approach can accurately capture the structure and characteristics of the supersonic combustion flow field.

Design of a high-resolution optical measuring ring for supersonic combustion flow field

Design of a high-resolution optical measuring ring for supersonic combustion flow field

一种提高超燃流场羟基示踪速度测量中信号提取能力的方法

一种提高超燃流场羟基示踪速度测量中信号提取能力的方法

The hybrid RANS/LES of partially premixed supersonic combustion using G/Z flamelet model

In order to describe partially premixed supersonic combustion numerically, G/Z flamelet model is developed and compared with finite rate model in hybrid RANS/LES simulation to study the strut-injection supersonic combustion flow field designed by the German Aerospace Center. A new temperature calculation method based on time-splitting method of total energy is introduced in G/Z flamelet model. Simulation results show that temperature predictions in partially premixed zone by G/Z flamelet model are more consistent with experiment than finite rate model. It is worth mentioning that low temperature reaction zone behind the strut is well reproduced. Other quantities such as average velocity and average velocity fluctuation obtained by developed G/Z flamelet model are also in good agreement with experiment. Besides, simulation results by G/Z flamelet also reveal the mechanism of partially premixed supersonic combustion by the analyses of the interaction between turbulent burning velocity and flow field.

Numerical Investigation of the Influence of Injection Scheme on the Ethylene Supersonic Combustion

Ethylene supersonic combustion flow field in different injection schemes was studied numerically in the flight Mach 4. The results show that injection pressure has significant influence on the location of the separation zone and the heat release region, but the starting point of the separation region was mostly influenced by the heat release rather than by the injection pressure; the combustion efficiency of the injection schemes including two injection points is higher than that of three injection points, while the total pressure recovery coefficient of the former injection schemes is lower than the latter; excessive ethylene injected in upstream will lead to the change of free-stream flow conditions, which behaves as the inlet unstart in practical application; more ethylene could be injected in downstream to avoid the problem; on the condition of avoiding thermal choke in isolator, it is more advantageous that injection points were arranged more closely to the starting point of separation zone in upside and to the front of the cavity in downside.

Flame stabilization in a supersonic combustor with hydrogen injection upstream of cavity flame holders: experiments and simulations

Experimental observations and numerical simulation were conducted to study the flame characteristics and flame stabilization mechanism in a supersonic combustor with hydrogen injection upstream of cavity flame holders. OH radical distribution of the reacting flowfield was obtained using OH spontaneous emission and OH-PLIF (planar laser-induced fluorescence). The supersonic combustion flowfield with L/ D = 7 cavity was calculated by large eddy simulation. The turbulence–combustion interaction model was based on a partially premixed flamelet model with a level-set approach. The results showed that hydrogen fuels were transported into the cavity shear; lean mixture and rich mixture were produced in the internal cavity and the declining jet, respectively. An approximately steady partially premixed flame front exists in the cavity shear layer. The flame front propagates and extends to the region around the fuel jet due to the interaction of counter-rotating vortices induced by the jet with the cavity shear layer. The flame front sustained in the shear layer likely penetrates the jet core and ignites the whole jet. Behind the flame front, most of the jet beam is burned as diffusion flames. The physical process of the flame stabilization demonstrated the similarity with triple flame characteristics, which indicted that triple flame theory might be the basic flame stabilization mechanism of the cavity flame holders.

Large Eddy Simulation of Hydrogen-Fueled Supersonic Combustion with Strut Injection

In order to obtain more accurate simulation results and properties of combustion in supersonic combustion flow fields, modules of large eddy simulation of reactive turbulent flow and fifth-order WENO scheme was developed. Large eddy simulation of hydrogen-fueled supersonic combustion with strut injection was conducted. Simulations results compare were with experimental measurements, which including wall pressure, velocity, velocity fluctuation and temperature.

Flame Characteristics in Supersonic Combustor with Hydrogen Injection Upstream of Cavity Flameholder

Flame characteristics and a plausible flameholding mechanism in a supersonic combustor, with hydrogen injection upstream of a cavity flameholder, were investigated in the present study. Instantaneous OH radical distribution of the combustion flowfield was obtained using OH planar laser-induced fluorescence. According to the similarity of experimental observations with different cavities, a typical L/D = 7 cavity was chosen, and its supersonic combustion flowfield with hydrogen injection was calculated by large-eddy simulation. The results showed that the cavity shear layer plays a very important role in the flameholding process. An approximately steady flame existed in the cavity shear layer and hot combustion products were transported into the injection jet by the vortex interaction ofthejet-with-cavity shear layer. Flame then spread gradually following the counter-rotating vortex induced by the jet until the whole injection jet was ignited. Combustion products, which generated from the cavity shear layer and the jet, were convected into the cavity by the unsteady motion of the cavity shear layer and transported with the recirculation flow to the cavity front wall. These hot products and their intermittent combustion then heated up the cavity, and the fuel that entered into the cavity shear layer was preheated. Thus, the flameholding cycle was formed.

Direct Measurements of Skin Friction in Supersonic Combustion Flow Fields

An experimental investigation was conducted to measure skin friction along the chamber walls of supersonic combustors. A direct force measurement device was used to measure simultaneously an axial and a transverse component of the small tangential shear force passing over a nonintrusive floating element. This measurement was made possible with a sensitive piezoresistive deflection sensing unit. The floating head is mounted to a stiff cantilever beam arrangement with deflection due to the flow on the order of 0.00254 mm (0.0001 in). This allowed the instrument to be a nonnulling type. A second gage was designed with active cooling of the floating sensor head to eliminate nonuniform temperature effects between the sensor head and the surrounding wall. The key to this device is the use of a quartz tube cantilever with piezoresistive strain gages bonded directly to its surface. A symmetric fluid flow was developed inside the quartz tube to provide cooling to the backside of the floating head. Tests showed that this flow did not influence the tangential force measurement. Measurements were made in three separate combustor test facilities. Tests at NASA Langley Research center consisted of a Mach 3.0 vitiated air flow with hydrogen fuel injection at Pt = 500 psia (3466 kPa) and Tt = 3000 R (1667 K). Two separate sets of tests were conducted at the General Applied Science Laboratory (GASL) in a scramjet combustor model with hydrogen fuel injection in vitiated air at Mach = 3.3, Pt = 800 psia (5510 kPa), and Tt = 4000 R (2222 K). Skin friction coefficients between 0.001–0.005 were measured dependent on the facility and measurement location. Analysis of the measurement uncertainties indicate an accuracy to within ± 10–15 percent of the streamwise component.

Parabolized Navier-Stokes analysis of ducted supersonic combustion problems

The extension of a recently developed implicit/explicit approach to analyzing two-dimensional ducted supersonic mixing problems with finite-rate chemistry is presented. The hybrid approach combines a fully implicit parabolic mixing algorithm, an explicit viscous-characteristic-based wave-solver algorithm, and a linearized implicit chemical kinetic solution algorithm. The resultant model provides spatial marching solutions of the parabolized Navier-Stokes (PNS) equations for supersonic combustion problems. Specialized procedures are incorporated to deal with the near-wall sublayer. A parabolic option is available that can be utilized in the direct or inverse mode for design applications. The ability of the model to treat shock waves and wave/mixing layer interactions is assessed via comparisons of predictions with those of well-establis hed, explicit shock-capturing Euler (SCIPPY) and PNS (SCIPVIS) models. Applications to several supersonic combustion flowfield problems are presented that exhibit overall capabilities.