Radiative heat transfer measurements of Titan atmospheric entry in a shock tube

Measurements were performed in the T6 Stalker facility operating in Aluminum Shock Tube mode for conditions relevant to Titan entry. Spatially and spectrally resolved radiation emitted from a high-temperature test gas behind a normal shock was recorded by means of emission spectroscopy. For Titan atmospheric entry, the main radiator of interest is cyano radical, formed in the nonequilibrium region behind the shock. The tests reported in this work measured radiation at velocities from 3.1 to 8.5 km/s and freestream pressures of 13, 20, and 133 Pa at a nominal composition of 98% N2 and 2% CH4. These shock layer radiation experiments employed an optical emission spectroscopy system on each side of the facility to allow two spectral regions to be measured simultaneously for each test, covering the spectral range of 200–900 nm. The present work provides a new, comprehensive benchmark set of data relevant to Titan entry studies

Proračun koeficijenta tlaka pri osnosimetričnom optjecanju rotacionog tijela metodom karakteristika

U radu je opisan postupak proračuna rasporeda tlaka na površini rotacionog tijela pri osnosimetričnom nadzvučnom optjecanju korištenjem metode karakteristika. Ova je metoda jedna od najučinkovitijih numeričkih metoda računalne dinamike fluida za rješavanje složenih problema optjecanja tijela. Promatran je slučaj dvodimenzijskog optjecanja vitkog tijela.Za rješavanje navedenog problema razvijen je program u računalnom paketu MATLAB, koji je testiran na primjeru iz literature za koji je poznata raspodjela tlaka. Unutar diplomskog rada opisani su algoritmi razvijeni u MATLAB-u pomoću kojih su izračunate točke mreže te potrebna svojstva toka fluida u njima. Nadalje, razvijeni je program upotrijebljen za određivanje raspodjele tlaka na približnom tijelu zrakoplova ''Concorde'' pri brzini krstarenja te nultom napadnom kutu. Naposljetku, izračunat je koeficijent valnog otpora tijela navedenog zrakoplova

Natural and Induced Transition on a 7deg Half-Cone at Mach 6

Pressure fluctuations caused by instabilities in the boundary layer are among the causes which lead to transition on a space vehicle during atmospheric re-entry. The knowledge of these unsteady fluctuations could help to identify the mechanisms which take part into the transition process and better predict them. In this framework and in support of the EXPERT PL4/PL5 post-flight analysis, surface pressure measurements have been performed in the VKI H3 Hypersonic Wind Tunnel. A 7deg half-angle cone with exchangeable nosetip was equipped with a stream-wise array of high frequency pressure transducers (PCB 132A31). Instabilities in the boundary layer have been investigated on a smooth surface and behind an isolated roughness element. Experimental data are then compared with linear stability theory computations, in support to the e^N transition prediction method and to the calibration of the ground facility. The results provided on a simplified ground test model, show the growth of second mode waves under a laminar boundary layer and their break down to turbulence as a function of Reynolds number and streamwise location. Moreover the effect of an isolated roughness element on the boundary layer has been characterized in terms of generated instabilities

Iron contamination in high-enthalpy test facilities: OH PLIF imaging considerations

Thermometry measurements were performed in a scramjet combustor using thermally-assisted laser-induced fluorescence. Experimental data were obtained through laser-induced fluorescence where a laser beam was focused into the combustor and the OH Q1(8) (v ''=0 -> v '=1) transition at 283.55 nm was excited. Measurements in the 360 nm spectral region resulted in detection of a strong fluorescence line. Investigation of this phenomenon led to the conclusion that when the OH Q1(8) line is chosen for excitation, another line of a different species that lies very close to Q1(8) is also unintentionally excited. The second species has been identified as iron, Fe I. The assumption is that the erosion of the shock tunnel walls introduced iron in the flow that is excited with the laser when exciting OH. Thus, it is very important to ensure that this is avoided in future. Instead, another OH excitation line should be used for OH excitation line in planar laser-induced fluorescence (PLIF) experiments where iron may potentially also be present in the flow. Alternatively, fluorescence from this iron line can be circumvented by proper filtering of the PLIF. A proposal is made to turn the iron contamination into advantage by using iron present in the non-combusting flow for PLIF

Thermal compression effects within a fundamental, hydrogen-fuelled scramjet

The impact of thermal compression on combustion has been studied experimentally in a hydrogen-fuelled, non-uniform scramjet flowpath. The experimental model consisted of a three-dimensional inlet, a constant area rectangular combustor, and a single-ramp expansion nozzle, all having a constant width, and was tested at an equivalent Mach 10 flight condition with 58 kPa dynamic pressure. Combustion was suppressed in different regions of the combustor by injecting helium as a replacement for hydrogen fuel in either spanwise half of the engine. Thermal compression effects increased the combustion-induced pressure rise by across the majority of the combustor for an equivalence ratio of 0.8, however a shock-related artefact dominated similar measurements for a case with an equivalence ratio of 1.0. Over a smaller region at the end of the combustor, the combustion-induced pressure rise was increased by and for equivalence ratios of 0.8 and 1.0, respectively. Time-integrated OH emission signals increased by 19% and 31%, whilst time-resolved signals increased by 54% and 35%, for equivalence ratios of 0.8 and 1.0, respectively. This study presents the first experimental evidence of thermal compression directly increasing combustion-induced pressure rise in a scramjet engine

Single-shot temperature measurements in a scramjet combustor using thermally assisted laser-induced fluorescence

The knowledge of the temperature in the flowfield of the scramjet is critical to the understanding of its performance. This study represents the application of an optical diagnostic technique, thermally assisted laser-induced fluorescence, to the temperature measurements in the combustor of a supersonic combustion engine. The experiments were conducted in the T4 shock tunnel using a scramjet model. Experimental data were obtained by focusing a laser beam into the combustor and exciting the OH radical. A detailed numerical model that simulates all radiative and collisional processes of relevance with the appropriate system of differential equations was developed for the analysis of the influence of energy transfer processes on the fluorescence signal. The population distributions achieved with numerical modeling of the LIF process validated the approach proposed to deduce temperatures. By evaluating the scramjet LIF spectra with a full spectral fit, a temperature distribution across the combustor width was obtained. The experimental results were compared with the computational fluid dynamics simulations of the combustion process

An experimental investigation of a thermal compression scramjet with OH imaging

Scramjet engines must be designed to operate over a range of Mach numbers. In order to maintain robust starting characteristics at low freestream Mach numbers, the inlet contraction ratio must be minimized, reducing performance at high Mach numbers. The technique of thermal compression can be used to maintain robust combustion at low mean pressure and temperature, improving the performance of low inlet contraction ratio engines over a range of Mach numbers. This paper examines a scramjet engine with a threedimensional inlet that induces spanwise gradients of pressure and temperature, producing high inlet compression on one side and low inlet compression on the other. The effect of combustion in each half on pressure throughout the entire flowfield is examined by suppressing combustion on either half of the combustor through injecting either a combusting fuel, hydrogen, or a non-combusting replacement, helium. Two optical techniques are used to examine how the concentration of the OH radical varies throughout the combustor. Images of OH* Chemiluminescence and OH Planar Laser-Induced Fluorescence (PLIF) map the concentration and production of OH radicals indicating chemical activity throughout the combustion process. The thermal compression effect is isolated, and is shown to increase combustion-induced pressure rise in a scramjet engine