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Purpose The purpose of this paper is to demonstrate the aerodynamic behavior of a supersonic combustion test bench (SCTB) components, as the transition piece and the combustor of a scramjet (supersonic combustion ramjet), manufactured by 3D printing or additive manufacturing (AM). Design/methodology/approach For the dimensional and structural analysis of the manufactured models, a portable 3D scanner was used to generate the mesh of its dimensions, and to compare them before and after the experiments, a roughness measuring system was also used to verify the roughness inside the models before and after the tests, as roughness is an important parameter because it directly affects the boundary layer. For the visualization of the flow, the non-intrusive schlieren optical technique was used. Findings The experiments were carried out on the SCBT for Mach 2 flows, using the manufactured prototypes and showed that there was no structural and dimensional change of the model after the test batteries. It was found that the roughness presented by the material did not affect the quality of the flow generated. This shows that the investigated material can also be applied in experiments with supersonic flow. Originality/value This paper presents that it is possible to use in ground test facilities, for the studies of supersonic flow (in cold condition), pieces and models manufactured by 3D printing without affecting the quality of the flow generated during the experiments. This study presents a new perspective to approach AM applied in the studies of supersonic flows.

Túneis de Choque são dispositivos capazes de gerar escoamentos de altas temperaturas e velocidades supersônicas, simulando as condições reais de escoamento na entrada de um combustor de motor Scramjet, viabilizando o estudo de combustão supersônica. Neste trabalho, são realizados ensaios aerotermodinâmicos e de combustão supersônica utilizando um Túnel de Choque, de modo a reproduzir as condições de escoamento na entrada de um combustor de motor Scramjet. O arranjo experimental consistiu na utilização do Túnel de Choque, um dispositivo com uma seção de alta pressão (Driver) e uma seção de baixa pressão (Driven), separadas entre si por um diafragma. Ao ser rompido o diafragma, ocorre a propagação do escoamento, o qual é acelerado a velocidades supersônicas através de um bocal bidimensional conectado à câmara de combustão supersônica. A metodologia adotada possibilitou a análise do escoamento no interior da câmara de combustão supersônica do motor Scramjet, gerando escoamentos de M=2.7.

This experimental campaign focused on the analysis of the feasibility for impulse generation of a repetitively pulsed airbreathing laser lightcraft at hypersonic speeds. The future application of interest for this basic research endeavor is the laser launch of nano- and microsatellites (i.e., 1–100 kg payloads) into low Earth orbit, at low cost and on demand. These laser propulsion experiments employed a hypersonic shock tunnel, integrated with twin gigawatt pulsed Lumonics 620-TEA CO2 lasers to produce the required test conditions. Time-dependent surface pressure distributions were recorded together with schlieren movies of the flowfield structure resulting from the laser energy deposition, at nominal Mach numbers ranging from 6 to 10. The laser induced breakdown occurring off-surface and across the inlet’s midchannel was analyzed and discussed against previous theoretical and experimental results. The data indicated laser induced pressure increases of 0.7–0.9 bar with laser pulse energies of ∼170 J, on ...

Resumen en: This paper presents a research in the development of the 14-X hypersonic airspace vehicle at Institute for Advanced Studies (IEAv) from Department of Sci...

Experimental results of bow shock wave mitigation by laser-plasma energy addition in a low-density Mach 7 hypersonic flow conducted in a shock tunnel are presented. A high-power pulsed CO2 laser operating with 7 J of energy and 30 MW of peak power was used to generate the plasma ahead of a hemispherical model installed in the tunnel test section. The schlieren technique was used to visualize the time evolution of energy addition to the flow by laser-induced plasma and the interaction between this disturbed region and the inherent bow shock formed on the model by hypersonic flow. A complete mitigation of the bow shock profile under action of the energy addition was observed. The impact pressure on the hemispherical model measured at the stagnation point reveals the correlation between the schlieren images and the pressure reduction.

The experimental results on the laser-supported directed-energy "air spike" (DEAS) in hypersonic flow are presented. A CO 2 TEA laser has been used in conjunction with the IEAv 0.3-m Hypersonic Shock Tunnel to demonstratethe laser-supported DEAS concept. A single laser pulse generated during the tunnel useful test time was focused through a NaCI lens ahead of an aluminum hemisphere-cylinder model fitted with a piezoelectric pressure transducer at the stagnation point. In the more recent experiments, a double Apollo disk model fitted with seven piezoelectric pressure transducers substituted the simple hemisphere-cylinder model. The objective of the present research is to corroborate the past results as well as to obtain additional pressure distribution information.

at a Ž eld point r 0, Er;d.r 0/, can be computed as E.r 0/ D E .r /DA .s/ei j ks (1) where E i .r/ is the Ž eld incident on the re ection or diffraction point r , D is a dyadic re ection or diffraction coefŽ cient,A.s/ is a spreading factor, and s is the distance from the re ection or diffraction point r to the Ž eld point r 0. D and A can be found from the geometry of the structure at re ection or diffraction point r and the properties of the incidentwave there. Conclusions The ISS is being assembled in space. Thus, conventional ground tests for communicationand navigationsystemperformanceare not possible with the ISS completely assembled. Rigorous computer simulations are important for ISS communication and navigation systemperformanceassessment.Both  ight data and computersimulation results presented in this Note indicate the solar array panels maybemodeledusingperfectre ectingplatesat S-bandfrequencies for communication and GPS signal strength predictions.

The Laser Propulsion (LP) experiments employed the T3 Hypersonic Shock Tunnel (HST), integrated with twin gigawatt pulsed Lumonics 620-TEA CO(2) lasers to produce the required test conditions. This effort was carried out at nominal Mach numbers ranging from 6 to 10. Time-dependent surface pressure distributions were recorded together with schlieren movies of the flow field structure resulting from the laser energy deposition. These visualizations of inlet and absorption chamber flowfields enabled the qualitative analysis of important phenomena impacting laser-propelled hypersonic air-breathing flight. The data indicated laser-induced pressure increases of 0.7- 0.9 bar with laser pulse energies of approx. 170 J, on off-shroud induced breakdown condition, and free stream Mach number of 6. The results of this research corroborate the feasibility of laser powered, airbreathing flight with infinite specific impulse (Isp=infinity).

Currently, a new generation of scientific aerospace vehicles, using advanced hypersonic airbreathing propulsion based on supersonic combustion technology, is in development at several research centers [1].

Today the access to space is done, only, by multi-stage rocket-powered vehicles, which have flown hypersonically, carrying their own propellant (solid and/or liquid, oxidizer along with fuel) to propel payloads and astronauts to Earth’s orbit.

The aerospace technological products have grown that one cannot conceive of putting payloads (satellites) into Earth orbit or beyond using technologies in operation (rockets carry out solid or liquid fuel). The knowledge required to keep the current launching vehicles is already so high that if the countries do not have a technological support for their own industry, they will depend on of the supplier countries and not have independent capacity sustained. Aerospace vehicle limitations for launching payloads into orbit or beyond require a continuous reduction in size, weight and power consumption of launch vehicles. Some solutions to these challenges require paradigm shifts, new production methods, and new technologies of strategic nature. The requirements of platformslaunched satellites, high performance and reliability, as well as the strict limitations of fuel (reduction of size, weight and power consumption) for launching payloads into orbit or beyond provide the development of hypersonic aircraft using hypersonic airbreathing propulsion based on supersonic combustion.

Flight-tests, computational fluid dynamic facilities and ground-based test facilities are the hypersonic methodologies used to define and to quantify the aerothermodynamic environments that exist during the fly of the aerospace vehicle into the Earth’s atmosphere at hypersonic flight speeds.

This article discusses the production of hypervelocitv -hypersonic flows in a combustion shock tunnel operating in the equilibrium interface mode. In this mode of operation, the additional compression provided by the approaching interface is used to obtain higher pressures and temperatures, as opposed to the reflected method. A computer code was developed to model the operation of a shock tunnel in the equilibrium interface condition. In this article, all the calculations were made for the Rensselaer Polytechnic Institute (RPI) 1.22-m-diam Combustion Driver Hypersonic Shock Tunnel. The major drawback of the interface compression technique, which is the contamination of the driven gas by the driver gas, was overcome through the utilization of a small volume region separating the two gases. Numerical results indicate that the RPI facility will be able to generate reservoir temperatures of the order of 20,000 K and reservoir pressures of the order of 30,000 psi. These reservoir conditions can be used to produce test section Mach numbers of 35.

Laser propulsion is an innovative concept of accessing the space easier and cheaper where the propulsive energy is beamed to the aerospace vehicle in flight from ground—or even satellite‐based high‐power laser sources. In order to be realistic about laser propulsion, the Institute for Advanced Studies of the Brazilian Air Force in cooperation with the United States Air Force and the Rensselaer Polytechnic Institute are seriously investigating its basic physics mechanisms and engineering aspects at the Henry T. Hamamatsu Laboratory of Hypersonic and Aerothermodynamics in Sao Jose dos Campos, Brazil. This paper describes in details the existing facilities and measuring systems such as high‐power laser devices, pulsed‐hypersonic wind tunnels and high‐speed flow visualization system currently utilized in the laboratory for experimentation on laser propulsion.

A three-dimensio nal axisymmeiric hypersonic engine inlet was investigated in the Rensselaer Polytechnic Institute Hypersonic Shock Tunnel at a Mach number range of 10-25 and at stagnation temperatures of 8004100 K. Static and total pressures were measured at various positions along the forebody and within the cowled engine to obtain estimates of inlet performance and data for validation of a computational fluid dynamic (CFD) code. Schlieren and luminosity photographs of the flowfield were obtained to aid the interpretation of pressure data. Results indicate significant viscous effects on the flow structure and resulting pressures. The Newtonian pressure distribution approximated the experimental results for the forebody well in general form, but the shock wave-boundary-layer interactions resulted in complex flow phenomena with pressure fluctuations on the vehicle surface that could not be predicted by the Newtonian theory. For ideal gas flows, the pressure recovery at the entrance of the heat addition region of the inlet was similar to that of an ideal 30 deg conical inlet.

A laser-propelled transatmospheric vehicle experiences the full spectrum of flight regimes, from subsonic to hypersonic, along its Earth-to-Orbit (ETO) trajectory. A 2-D model of a laser propelled LightCraft was tested in the T3 Hypersonic Shock Tunnel (HST) at the Henry T. Nagamatsu Laboratory for Aerothermodynamics and Hypersonics. The test campaign began with no flow, ambient (1 bar) conditions, and transitioned to medium enthalpy HST runs at Mach number of 9.0. The model was instrumented with piezoelectric pressure transducers to measure the time dependent pressure distributions. A high speed digital Schlieren imaging system recorded the time-dependent flowfield evolution inside the LightCraft absorption chamber, following pulsed laser energy deposition. This data was also compared with blast wave arrival at numerous pressure transducers distributed along the 2D engine centerline. A Lumonics TEA 622 laser system with an unstable resonator cavity emitted 1 μs, 200 J pulses (low divergence, 90ns FWHM) into the 0.6 m diameter HST test section.

The interaction between a propulsive laser beam and the flowfield after the shock wave, generated during the flight of a beam-powered conical-shaped transatmospheric vehicle, is studied. Although the adopted theoretical model is a simplified one, the results of this study anticipate some of the aero-optical phenomena to be expected during a beamed transatmospheric flight. By considering a propulsive beam operating at 3.8 /an, ohase distortions up to 51.41 wavelengths are reported. The range of flight Mach numbers covered by the present analysis falls between 1.67 and 20. As a major conclusion, considerable beam degradation can be expected only during the beginning of the beamed transatmospheric flight.

A new 0.6‐m. diameter Hypersonic Shock Tunnel is been designed, fabricated and will be installed at the Laboratory of Aerothermodynamics and Hypersonics IEAv‐CTA, Brazil. The brand new hypersonic facility, designated as T3, is primarily intended to be used as an important tool in the investigation of supersonic combustion management and of electromagnetic energy addition for flow control. The design of the runnel enables relatively long test times, 2–10 milliseconds, suitable for basic supersonic combustion and energy addition by laser experiments. Free stream Mach numbers ranging from 6 to 25 can be produced and stagnation pressures and temperatures of 200 atm. and 5,500 K, respectively, can be generated. Shadowgraph and schlieren optical techniques will be used for flow visualization and the new facility is expected to be commissioned by the end of 2006.

Experimental results on the visualization of the time evolution of the laser‐plasma induced breakdown produced in low density hypersonic flow using the Schlieren technique are presented. The plasma was generated by focusing the high power laser pulse of a CO2 TEA laser in the test section of the IEAv 0.3m Hypersonic Shock Tunnel. An ultra‐high speed electronic tube camera was used to register the event. The photographs reveal the expansion of the shock wave produced by the laser generated hot plasma and the convection of the plasma kernel by the hypersonic flow. It is also observed the interaction between the plasma disturbed region and the shock established by the flow around an hemisphere‐cylinder model. A strong change in the shock wave structure near the model was observed, corroborating the DEAS concept.

Flow visualization of laser energy addition upstream blunt bodies in hypersonic flow was achieved by means of open shutter photography and high‐speed cameras techniques. The laser energy was supplied by two CO2 TEA lasers, sharing the same optical cavity, while the hypersonic flow was generated in the IEAv 0.3m Hypersonic Shock Tunnel — HST test section. Single and double laser pulses, generated during the tunnel useful test time, were focused through a NaCl lens upstream a simple hemisphere‐cylinder and Double Apollo Disc models. A regular SLR camera, a high‐speed CCD camera and an ultra‐high speed electronic tube camera provided imaging. Preliminary results using the ultra‐high speed camera in a shadowgraph/schlieren arrangement revealed the dynamic complex flow structure that is established during the addition of the laser energy to the hypersonic flow upstream a blunt body.

The present paper presents recent (original) and previous experimental results on the Laser‐Supported Directed Energy “Air Spike” — DEAS in hypersonic flow achieved by the Laboratory of Aerothermodynamics and Hypersonics — LAH, Brazil. A CO2 TEA laser has been used in conjunction with the IEAv 0.3m Hypersonic Shock Tunnel — HST to demonstrate the Laser‐Supported DEAS concept. A single laser pulse generated during the tunnel useful test time was focused through a NaCl lens ahead of an aluminum hemisphere‐cylinder model fitted with a piezoelectric pressure transducer at the stagnation point. In the more recent experiments, the simple hemisphere‐cylinder model was substituted by a Double Apollo Disc model fitted with seven piezoelectric pressure transducers and six platinum thin film heat transfer gauges. The objective being to corroborate the past results as well as to obtain additional pressure distribution information and new heat transfer data.

Aerodynamic shear loading tests were conducted on two different designs of the Shroud Deployment System (SDS) bladder in the RPI 24-m diameter High Pressure Shock Tube. The objective of this test was to determine the maximum aerodynamic shear load the shroud ejection bladders (series -301 and -303) could withstand before being sheared off the substructure. The bladders were subjected to impulsive loading by applying various strength shock waves. The strength of the shock waves were varied by changing the ratio of driven and driver tube pressures. The deployment bladder was mounted at the exit of the shock tube in the 200-fi3 dump tank on a double strut confIguration. The maximum total pressures at failure were 404 psi for model -301 and 475 psi for model -303 f5.25%.

STRACT’ An investigation ‘of a Shroud Deployment System (SDS) bladder was conducted in the Rensselaer Polytechnic Institute 2+in .Hypersonic Shock Tunnel. The cut-rem investigation experimentally determined the axial drag coefficient for the bladder, while attached ~0 .$he SDS d-ge support, for nominal Mach numbers of 8 to 20 and stagnation temperatures up to 7400 “R Me&ements were made of the pitot pressure and the force applied to the bag by the j~ypynjc: ,flqw. i 10..