Fluid Mode Excitation in Launch Vehicle Feed Lines Induced by Pogo Accumulator Venting

2014 ◽  
Author(s):  
Kirk W. Dotson ◽  
Brian H. Sako ◽  
Trinh T. Nguyen
Keyword(s):  
Space Vehicle ◽  
Launch Vehicle ◽  
Liquid Oxygen ◽  
Rocket Engines ◽  
Launch Vehicles ◽  
Feed Line ◽  
Large Pressure ◽  
Structural Responses ◽  
Flow Oscillations ◽  
Liquid Rocket

Launch vehicles with liquid rocket engines have feed lines through which propellants flow to the engine. To prevent feedback between structural responses and propellant pressure and flow oscillations, a compliant device called a pogo accumulator is typically installed in the propellant feed line. Even if a catastrophic interaction is thus averted, the fluid-induced structural responses may exceed those for important flight events such as liftoff and atmospheric buffeting. In that case, the fluid-induced excitation must be predicted in order to ensure adequate structural margins for the launch vehicle and space vehicle hardware. Venting of compliant gas in the pogo accumulator prior to engine shutdown is known to exacerbate the fluid-induced excitation. In particular, for the Atlas V launch vehicle, a 5–7 Hz fluid mode with large pressure gains at the aft end of the liquid oxygen feed line often excites structural modes just prior to engine cutoff. A methodology for the prediction of these structural responses is presented.

scholarly journals Development of Solid Gas Generating Compositions to Ensure Non Explosiveness of Spent Orbital Stages of Liquid Rocket of Space Launch Vehicles

2017 ◽  
Vol 19 (1) ◽  
pp. 63
Author(s):  
V. Trushlyakov ◽  
K. Zharikov ◽  
D. Lempert
Keyword(s):  
Rocket Engine ◽  
Drainage System ◽  
Combustion Products ◽  
Launch Vehicle ◽  
Liquid Oxygen ◽  
Launch Vehicles ◽  
Gas Generator ◽  
Space Launch ◽  
Surrounding Space ◽  
Liquid Rocket

The choice is discussed of solid gas generating compositions for venting by hot combustion products a fuel tank of the spent orbital stage of a space launch vehicle with a main liquid rocket engine. Non explosiveness is achieved via eliminating the<br />possibility of freezing the drainage system when products of gasification (vapours of a propellant component + the remains of a gas boost + the hot products of combustion of solid gas generating compositions) are discharged from the tank into surrounding space. There are imposed requirements, constraints, and criteria for selecting solid gas generating compositions. When considering tank with the residues of liquid oxygen belonging to orbital spent stage of the launch vehicle «Zenith» the ways are shown how to ensure explosion safety, which on the basis of proposed approaches by selecting solid gas generating compositions (SGC) which generate oxygen and<br />nitrogen. As a criterion of choice of SGC the total mass of the gasification system is adopted, which includes the SGC mass for gasification of liquid propellant residues, the mass of the gas generator and the mass of system to supply the combustion products of SGC into the tank. It is proposed use of residual heat in the condensed phase of the SGC combustion products to heat up the drainage system, which will increase the probability of a trouble-free operation of the drainage system.

Interaction of Liquid Propellant and Coupled System Structure for Atlas V Launch Vehicle

2010 ◽  
Author(s):  
Kirk W. Dotson ◽  
Brian H. Sako ◽  
Daniel R. Morgenthaler
Keyword(s):  
Hydraulic System ◽  
Coupled System ◽  
Launch Vehicle ◽  
System Structure ◽  
Liquid Oxygen ◽  
Launch Vehicles ◽  
Liquid Propellant ◽  
Pressure Oscillations ◽  
Maximum Acceleration ◽  
Structure Interaction

In structural modeling of launch vehicles, liquid propellant is sometimes rigidly attached to feedline walls. This assumption precludes the interaction of structural modes with propellant pressure and flow. An analysis of fluid-structure interaction (FSI) for the Atlas V launch vehicle revealed that structural models with rigidly-attached propellant yield unconservative response predictions under some conditions. In particular, during the maximum acceleration time of flight, pressure oscillations acting at bends in the Atlas V liquid oxygen (LO2) feedline excite 15–20 Hz structural modes that have considerable gain on the feedline and at the spacecraft interface. The investigation also revealed that the venting of gas from the pogo accumulator is an excitation source and changes the dynamic characteristics of the hydraulic system. The FSI simulation produced during the investigation can be adapted to mission-specific conditions, such that responses and loads are conservatively predicted for any Atlas V flight.

scholarly journals Hot-fire testing of liquid oxygen/hydrogen single coaxial injector at high-pressure conditions with optical diagnostics

2019 ◽  
Author(s):  
D.I. Suslov ◽  
J.S. Hardi ◽  
B. Knapp ◽  
M. Oschwald
Keyword(s):  
High Pressure ◽  
High Speed ◽  
Optical Diagnostics ◽  
Liquid Oxygen ◽  
Experimental Investigations ◽  
Rocket Engines ◽  
Chamber Length ◽  
Substantial Progress ◽  
Liquid Rocket ◽  
Coaxial Injector

Injector behavior is of utmost importance for the performance and stability of liquid rocket engines (LREs). A major problem is getting a highly efficient homogeneous mixture and effective chemical reaction of fuels at minimum chamber length. Despite substantial progress in numerical simulations, a need for experimental data at representative conditions for development and validation of numerical design tools still exists. Therefore, in the framework of the DLR-project “ProTau,” the authors have performed tests to create an extended data base for numerical tool validation for high-pressure liquid oxygen (LOx) / hydrogen combustion. During the experimental investigations, a windowed DLR subscale thrust chamber model “C” (designated BKC) has been operated over a broad range of conditions at reduced pressures of approximately 0.8 (4 MPa), 1 (5 MPa), and 1.2 (6 MPa) with respect to the thermodynamic critical pressure of oxygen. Liquid oxygen and gaseous hydrogen (GH2) have been injected through a single coaxial injector element at temperatures of ~ 120 and ~ 130 K, respectively. High-speed optical diagnostics have been implemented, including imaging of OH* emission and shadowgraph imaging at frequencies from 8 up to 10 kHz to visualize the flow field.

Increasing the specific impulse of liquid oxygen + kerosene rocket engines by introducing hydrogen into the combustion chamber

2019 ◽  
Author(s):  
S.A. Orlin ◽  
A.V. Orlov
Keyword(s):  
Combustion Chamber ◽  
Thermodynamic Analysis ◽  
Specific Impulse ◽  
Launch Vehicle ◽  
Liquid Oxygen ◽  
Rocket Engines ◽  
State Technical ◽  
Analytical Studies ◽  
Technical University ◽  
Engine Type

The investigation carried out at the Bauman Moscow State Technical University is aimed at establishing whether it may be possible to increase the specific impulse of liquid oxygen + kerosene rocket engines. It involved analytical studies of increasing specific impulse by introducing hydrogen into the oxygen/kerosene propellant. We confirm that, in the case of the oxygen/kerosene propellant used in the first stage engines, introducing hydrogen into the combustion chamber may increase its specific impulse. The results of our thermodynamic analysis show that the specific impulse increase is a function of the mass of hydrogen introduced. This enables the same engine type to be used for the first and second stages of a launch vehicle, which makes the whole system considerably less expensive and more reliable.

scholarly journals Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications

2012 ◽  
Vol 2012 ◽  
pp. 1-31 ◽  
Author(s):  
Bruce Chehroudi
Keyword(s):  
High Pressure ◽  
Critical Point ◽  
Space Shuttle ◽  
Specific Impulse ◽  
Liquid Oxygen ◽  
Liquid Jets ◽  
Rocket Engines ◽  
Liquid Rocket Engines ◽  
Liquid Rocket ◽  
Air Force Research Laboratory

Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME) or Vulcain (Ariane 5), the injected liquid oxygen finds itself in a supercritical condition. Very little detailed information was available on the behavior of liquid jets under such a harsh environment nearly two decades ago. The author had the opportunity to be intimately involved in the evolutionary understanding of injection processes at the Air Force Research Laboratory (AFRL), spanning sub- to supercritical conditions during this period. The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem.

Technology developments for thrust chambers of future launch vehicle liquid rocket engines

2003 ◽  
Vol 53 (4-10) ◽  
pp. 597-605 ◽  
Author(s):  
H. Immich ◽  
J. Alting ◽  
J. Kretschmer ◽  
D. Preclik
Keyword(s):  
Launch Vehicle ◽  
Rocket Engines ◽  
Liquid Rocket Engines ◽  
Liquid Rocket

Ecological aspect of launch vehicles development by criterion of minimal cost

2014 ◽  
Vol 25 (3-4) ◽  
pp. 114-119
Author(s):  
A. A. Baldin
Keyword(s):  
Space Debris ◽  
Low Cost ◽  
Launch Vehicle ◽  
Rocket Engines ◽  
Launch Vehicles ◽  
Rocket Propulsion ◽  
Second Stage ◽  
Maximal Performance ◽  
Stage Separation ◽  
The Future

One of the topical problems in modern aerospace engineering is accordance between ecological requirements and performance of the vehicle. On the other hand, problem of economical efficiency leads to change of the main criterion of designing to the minimization of costs (instead of maximal performance). According to modern trends of “low-cost” vehicles, different concepts of the future cost-effective launch vehicles are considered. It is necessary to validate these concepts according to requirements of ecological safety for the purpose of detection of the dominant launch vehicle configuration. Typical configurations of the future 'low-cost' launch vehicle are presented by 6 conceptual groups (Koelle, 2001). Conceptual group 1 (CG1) is presented by the Ballistic “Single stage to orbit” (SSTO) reusable vehicle. All vehicles which use classical rocketry scheme of the propulsion trajectory are called “Ballistic” i.e. the ballistic vehicle is lifted to orbit under the impact of rocket engines thrust. CG1-vehicle is able to reach the low earth orbit (LEO) without stage separation reducing the number of required rocket engines. Technological feasibility of SSTO concepts is proven by numerous studies (Koelle, 2001). CG2 representatives are ballistic “Two stages to orbit” (TSTO) reusable vehicles. The difference between CG1 and CG2 consists in application of vacuum rocket engines in the second stage  and, consequently, stage separation. CG2 are the most mass-effective vehicles. CG3 is presented by the winged SSTO vehicles with rocket propulsion by “Lifting body” aerodynamic scheme. Ascensional force is provided by the aerodynamic shape of the vehicle’s structure at high speeds. Winged TSTO vehicles with rocket propulsion and parallel or tandem staging form the CG4. The winged configuration provides wide landing capability for both stages. CG5 is presented by winged TSTO vehicles with airbreathing propulsion in the first stage and rocket-propelled second stage. Airbreathing jet engines provide high reusability ratio comparing with other concepts as well as the widest landing capability. Aerospace Plane with scramjet-rocket propulsion forms CG6. The vehicle is able to reach near-cosmic speed in rarefied layers of the atmosphere and then accelerate with rocket engines. The most ecologically important resemblance of represented concepts is reusability. This reduces space debris formation (due to lack of waste hardware). Reusable launch vehicles can also be used to return the spent satellites. Structural differences between the concepts form 3 criterions of comparison by ecological impact: 1) propellant toxicity; 2) safety of surface facilities (vehicle damage inside the atmosphere); 3) probability of space debris formation (vehicle damage outside the atmosphere). Comparison of the concepts by these criterions allows substantiating the most ecologically acceptable direction of research. Results of the comparison demonstrate that the most ecologically acceptable low-cost launch vehicle configuration is: Ballistic SSTO or TSTO reusable launch vehicle with “LOX+LH2” propellant. The results can be explained by following way: combustion products of the propellant “liquid oxygen + liquid hydrogen” are absolutely safe for environment. It also provides maximal performance of rocket engine (due to the highest specific impulse). Ballistic ascent scheme allows using relatively simple technologies and provides high reliability level. In combination with minimal time of atmospheric flight this provides high level of safety for surface facilities. These results may be used for substantiation of dominant research direction.

scholarly journals Numerical Investigation on the Thermal Behaviour of a LOx/LCH4 Demonstrator Cooling System

Aerospace ◽  
2021 ◽  
Vol 8 (6) ◽  
pp. 151
Author(s):  
Daniele Ricci ◽  
Francesco Battista ◽  
Manrico Fragiacomo
Keyword(s):  
Thermal Behaviour ◽  
Cooling System ◽  
Thermal Analyses ◽  
Operating Conditions ◽  
Liquid Oxygen ◽  
Rocket Engines ◽  
Successful Operation ◽  
Combustion Gases ◽  
Cooling Jacket ◽  
Liquid Rocket

Reliability of liquid rocket engines is strictly connected with the successful operation of cooling jackets, able to sustain the impressive operative conditions in terms of huge thermal and mechanical loads, generated in thrust chambers. Cryogenic fuels, like methane or hydrogen, are often used as coolants and they may behave as transcritical fluids flowing in the jackets: after injection in a liquid state, a phase pseudo-change occurs along the chamber because of the heat released by combustion gases and coolants exiting as a vapour. Thus, in the development of such subsystems, important issues are focused on numerical methodologies adopted to simulate the fluid thermal behaviour inside the jackets, design procedures as well as manufacturing and technological process topics. The present paper includes the numerical thermal analyses regarding the cooling jacket belonging to the liquid oxygen/liquid methane demonstrator, realized in the framework of the HYPROB (HYdrocarbon PROpulsion test Bench) program. Numerical results considering the nominal operating conditions of cooling jackets in the methane-fuelled mode and the water-fed one are included in the case of the application of electrodeposition process for manufacturing. A comparison with a similar cooling jacket, realized through the conventional brazing process, is addressed to underline the benefits of the application of electrodeposition technology.

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