The conceptual design of a hydrogen/ethylene fueled detonation ramjet (DR) studied earlier is updated based on numerical simulations and test fires at airflow Mach numbers 1.5, 2.0, and 2.5. The goal of the design update is to increase the operation stability, reduce aerodynamic drag, and improve thrust performance at continuous detonation of hydrogen/ethylene in the DR combustor. The updated DR is fabricated and tested. The thrust performance of the updated DR is increased significantly. At on-design M=2.0, the range of stable operation of hydrogen fueled DR combustor is expanded in terms of the air-to-fuel equivalence ratio from 1.6 to 3.3. At the air-to-fuel equivalence ratio of 3.1, the fuel-based specific impulse reached 4760 s. The maximum value of total thrust is 1000 N. At off-design M=1.5 and 2.5, the fuel-based specific impulse / total thrust attained 2780 s / 740 N, and 3780 s / 1160 N, respectively. Replacement of hydrogen by ethylene leads to the significant narrowing of the DR stable operation domain: continuous-detonation of ethylene–air mixture is registered only for near-stoichiometric mixtures.

The conceptual design of the hydrogen-fueled detonation ramjet (DR) of a new type for a cruising flight speed of Mach 2 at sea level is developed using 3D numerical simulations of the operation process. The calculated effective thrust of such a DR is shown to become positive at M = 1.3, i.e., the startup Mach number for such a DR can be lower than M ≈ 2.0, which is typical for ramjets operating on continuous-deflagration combustion. A DR demonstrator is designed and manufactured. Its test fires are performed in a blowdown wind tunnel (WT) at free air jet Mach numbers M = 2.0, 1.5, and 0.9. The result of test fires is the experimental proof of the possibility of arranging stable continuous-detonation combustion of hydrogen in the DR of the developed design at both Mach numbers exceeding 1.0. The maximum measured values of the fuel-based specific impulse and total thrust were 1610 s and 650 N for the tests with M = 1.5 and 1630 s and 860 N for the tests with M = 2.0.

The conceptual design of a hydrogen-fueled detonation ramjet (DR) of a new type for a cruising flight speed of Mach 2 at sea level is developed using multivariant three-dimensional (3D) numerical simulations of the operation process. The possibility of arranging the continuous-detonation (rotating or longitudinally pulsating) combustion of hydrogen in an expanding annular combustor of the DR is proved for the first time. 3D numerical simulations of the operation process in the DR for flight conditions with a Mach number M ranging from 1.1 to 2.7 are performed. The calculated effective thrust of such a DR is shown to become positive at M = 1.3, i.e., the startup Mach number for such a DR can be lower than M = 2.0, which is typical for ramjets operating on continuous-deflagration combustion. A DR demonstrator is designed and manufactured. Its test fires are performed in a pulsed wind tunnel (WT) at free air jet Mach numbers M = 2.0 and M = 1.5. The most important result of test fires is the experimental proof of the possibility of arranging stable continuous-detonation combustion of hydrogen in the DR of the developed design at both Mach numbers.

The conceptual design of a hydrogen-fueled dual-duct detonation ramjet (DR) of a new type for a cruising flight speed of Mach 2 at sea level is developed using multivariant three-dimensional (3D) numerical simulations of the operation process. The possibility of arranging the continuous-detonation (rotating or longitudinally pulsating) combustion of hydrogen in an expanding annular combustor of the DR is proved for the first time. The calculated effective thrust of such a DR is shown to become positive at M = 1.3, i. e., the startup Mach number for such a DR can be lower than M = 2.0 which is typical for ramjets operating on continuous-deflagration combustion. A DR demonstrator is designed and manufactured. Its test fires are performed in a pulsed wind tunnel (WT) at free air jet Mach numbers M = 2.0 and 1.5. The most important result of test fires is the experimental proof of the possibility of arranging stable continuous-detonation combustion of hydrogen in the DR of the developed design at both Mach numbers.

A mathematical model for simulating combustion and detonation of a fuel–air mixture in the gas cavity above the free water surface is developed. The model is based on solving the conservation equations of mass, momentum, and energy for a two-phase reacting gas–water medium with the phases treated as interacting interpenetrating continua having their own values of velocity, temperature, and turbulence characteristics. The model is validated by laboratory experiments. The test rig included a transparent cylindrical tube with one closed-end, a pool with an optically transparent window, as well as power, ignition, control, and measurement systems. The tube was vertically immersed with its open end in water and filled with a gaseous explosive mixture. In the experiments, a stoichiometric propane–air mixture was ignited and burned in the semi-closed 60 mL cylindrical volume above the free surface of water. The model is shown to predict satisfactorily the lift force acting on the tube, the time history of pressure in the volume, and the dynamics of the flame and gas–water interface motion during combustion in the volume. The model is intended to be applied for the design of boats with propulsion solely by combustion/detonation of fuel–air mixture in cavities constructed into a bottom surface of the boat. This propulsion system replaces conventional propellers, thereby reducing hydrodynamic resistance.

For reducing the hydrodynamic drag of a boat, a gas cavity can be made under the boat bottom, which will partially isolate the bot- tom from direct contact with water and provide ¤gas lubrication¥ by forced supply of atmospheric air or exhaust gases from a boat motor.

Выполнены расчетно-экспериментальные исследования по доводке конструкции макета-демонстратора (МД) детонационного прямоточного воздушно-реактивного двигателя (ДПВРД), использующего водород в качестве горючего. Доводочные работы с применением вычислительной технологии ФИЦ ХФ РАН были направлены на повышение устойчивости воздухозаборного устройства (ВЗУ) при М = 2,0, снижение аэродинамического сопротивления и повышение тяговых характеристик МД ДПВРД. Изготовлен новый вариант МД и проведены его огневые испытания в импульсной аэродинамической трубе (АТ) при числах Маха набегающего воздушного потока М = 1,5, 2,0 и 2,5. Важнейший результат огневых испытаний-значительное повышение тяговых характеристик нового варианта МД по сравнению со старым вариантом МД. Так, при M = 1,5 увеличение полной тяги и удельного импульса по топливу достигало 200 Н и 1100 с соответственно, а при М = 2,0-400 Н и 1300 с соответственно. Кроме того, при М = 2,0 значительно расширена область устойчивой работы камеры сгорания (КС) с непрерывно-детонационным горением водорода: от значения коэффициента избытка воздуха (КИВ) ~ 1,6 до ~ 3,3, причем при КИВ = 3,1 значение удельного импульса достигало ~ 4760 с. При M = 2,5 в ДПВРД данного типа впервые получен устойчивый непрерывно-детонационный рабочий процесс. Максимальные значения полной тяги и удельного импульса в испытаниях с M = 2,5 составили 1160 Н и 3780 с соответственно.

To reduce the hydrodynamic drag force to the movement of the boat, an artificial gas cavity is organized under its bottom. Such a cavity partially insulates the bottom from direct contact with water and provides “gas lubrication” by means of forced supply of atmospheric air or exhaust gases from the main propulsion system. A proper longitudinal and transverse shaping of the gas cavity can significantly (by 20%-30%) reduce the hydrodynamic drag of the boat at low (less than 3%) consumption of the propulsion system power for gas supply.

Modern high-speed unmanned aerial vehicles are powered with small-size turbojets or ramjets. Existing ramjets operating on the thermodynamic cycle with de§agrative combustion of fuel at constant pressure are efficient at flight Mach numbers M ranging from about 2 to 6.

The air-breathing pulsed detonation engine (PDE) for an aircraft designed for a subsonic flight when operating on the products of pyrolysis of polypropylene was developed using the analytical estimates and parametric multivariant threedimensional (3D) calculations. The PDE consists of an air intake with a check valve, a fuel supply system, a prechamber-jet ignition system, and a combustion chamber with an attached detonation tube. Parametric 3D calculations allowed choosing the best length of the PDE combustor, which provides an efficient mixing of air with fuel, the best way to ignite the mixture (prechamber-jet ignition), the best location of the prechamber, the minimum length of the section with turbulizing obstacles for flame acceleration and deflagration-to-detonation transition (DDT), and the best degree of filling the detonation tube with the fuel mixture to achieve the maximum completeness of combustion.

The thrust module (TM) for an aircraft designed for a subsonic §ight at a speed of 30 to 120 m/s when operating on a standard aviation kerosene TS-1 was developed using the analytical estimates and parametric multivariant three-dimensional (3D) calculations. The TM consists of an air intake with a check valve, a fuel supply system, a prechamber-jet ignition system, and a combustion chamber with an attached detonation tube. An experimental sample of TM was fabricated and its ¦re tests were carried out on a test rig with a thrustmeasuring table. In ¦re tests, TM characteristics are obtained in the form of dependencies of e ̈ective thrust, aerodynamic drag, and fuel-based speci¦c impulse on fuel consumption at di ̈erent speeds of the approaching air §ow. It has been experimentally shown that the fuel-based speci¦c impulse of the TM reaches 1000€1200 s, and the e ̈ective thrust developed by it reaches 180€200 N. The results of catapult launching tests of an unmanned aerial vehicle (UAV) powered with one and two paired TMs are also presented. The autonomous §ight of the UAV with a new type of power plant is demonstarted.

Проведены экспериментальные исследования процессов пульсирующего горения пропано-воздушной смеси в модельной днищевой каверне судна (без обводов судна), погруженной в бассейн с покоящейся водой. В экспериментах регистрировались расходы воздуха и горючего, распространение пламени, а также выталкивающая и толкающая силы, действующие на модельную каверну. Результаты экспериментов сравниваются с результатами трехмерных расчетов, основанных на физико-математической модели горения подготовленной горючей смеси в полузамкнутом объеме над свободной поверхностью воды, разработанной ранее: по форме и положению фронта пламени и границы раздела сред«газ-вода» в разные моменты времени и по динамике изменения сил, действующих на днище и на редан каверны. Получено удовлетворительное качественное и количественное согласие результатов расчетов и измерений.

The air-breathing pulsed detonation thrust module (TM) for an aircraft designed for a subsonic flight at a speed of up to 120 m/s when operating on a standard aviation kerosene was developed using the analytical estimates and parametric multivariant three-dimensional (3D) calculations. The TM consists of an air intake with a check valve, a fuel supply system, a prechamber-jet ignition system and a combustion chamber with an attached detonation tube. An experimental sample of TM was fabricated, and its firing tests were carried out on a test rig with a thrust-measuring table. In firing tests, TM characteristics are obtained in the form of dependencies of effective thrust, aerodynamic drag and fuel-based specific impulse on fuel consumption at different speeds of the approaching air flow. It has been experimentally shown that the fuel-based specific impulse of the TM reaches 1000-1200 s, and the effective thrust developed by it reaches 180–200 N.

Multidimensional calculations are performed to demonstrate that, by its characteristics, the pulse detonation engine (PDE) is a unique type of ramjet propulsion system, which can be used in both subsonic and supersonic aircraft. By a number of examples, it is shown that, in various thrust characteristics, such as the specific impulse, specific fuel consumption, and specific thrust, the PDE substantially exceeds ramjet engines.