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Thrust Augmentation Nozzle (TAN) Concept for Rocket Engine Booster Applications Scott Forde∗, Mel Bulman, Todd Neill

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Thrust Augmentation Nozzle (TAN) Concept for Rocket Engine Booster Applications Scott Forde∗, Mel Bulman, Todd Neill Acta Astronautica 59 (2006) 271–277 www.elsevier.com/locate/actaastro Thrust augmentation nozzle (TAN) concept for rocket engine booster applications Scott Forde∗, Mel Bulman, Todd Neill Aerojet, Sacramento, CA, USA Abstract Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results. © 2006 Published by Elsevier Ltd. 1. Introduction is less efficient in producing thrust. This is due to the gases over-expanding to a pressure below ambient. This Conventional rocket engines for a launch vehicle results in a portion of the nozzle generating negative booster stage need to deliver high thrust when taking thrust. At extreme area ratios, the exhaust jet will sep- off with the greatest vehicle weight, typically near arate from the nozzle, causing large transient loads and sea-level operation. They then operate until reaching high local heat fluxes, potentially damaging the noz- altitudes that have relatively low ambient pressures zle. A variable area nozzle would add complexity, cost, around the engine. weight, and size to the engine while still yielding less The vehicle requires engines with as high specific thrust at sea level than at vacuum. impulse (Isp) as practical to minimize propellant mass; however, a high-vacuum Isp engine requires a large area 2. TAN description ratio nozzle. These two requirements conflict since the large area ratio nozzle operating at sea-level pressure Aerojet’s patented TAN concept [1], shown in Fig. 1, overcomes these conventional engine limitations by in- jecting propellants and combusting in an annular region ∗ within the divergent section of the nozzle. This injection Corresponding author. E-mail addresses: [email protected] (S. Forde), of propellants at moderate pressures allows for obtain- [email protected] (M. Bulman), [email protected] ing high thrust at takeoff without overexpansion thrust (T. Neill). losses. The main chamber is operated at a constant 0094-5765/$ - see front matter © 2006 Published by Elsevier Ltd. doi:10.1016/j.actaastro.2006.02.052 272 S. Forde et al. / Acta Astronautica 59 (2006) 271–277 Nomenclature C∗ characteristic velocity TPA turbopump assembly Isp specific impulse (lbf-s/lbm) TVC thrust vector control LOX liquid oxygen T/W engine thrust per weight RP-1 rocket engine grade kerosene nozzle area ratio (nozzle exit TAN thrust augmented nozzle diameter/combustion chamber TCA thrust chamber assembly throat diameter) pressure while maintaining a constant head rise and The propellant combinations that have been tested in- flow rate of the main propellant pumps. Engine thrust clude gaseous oxygen (GOX) and LOX for the oxidizer augmentation greater than 100% of a normal engine is with hydrogen and RP-1 for the fuels. All other bipro- achievable. pellant combinations seem to be feasible with the TAN The concept is an extension of the liquid oxygen concept. This includes options of using propellant for (LOX)-augmented nuclear thermal rocket (LANTR) the TAN section that are different than the propellant [2–6], where LOX was injected into the divergent noz- used for the core engine. One such tripropellant option zle section of the superheated liquid hydrogen (LH2) uses LOX for both the main engine and the TAN oxi- exhaust of a nuclear thermal rocket thrust chamber dizer with hydrogen for the fuel on the core engine and assembly (TCA) to combust with the hydrogen and a heavier hydrocarbon fuel for the TAN injector. generate additional thrust. TAN takes the next step The TAN propellants can be supplied based on the and injects both oxidizer and fuel into the divergent optimum vehicle configuration. The three approaches nozzle section of the TCA of a conventional bipro- that have been evaluated are supplying propellants pellant booster rocket engine where the secondary from modified core engine boost pumps, incorporat- propellants mix and combust in the nozzle. This re- ing TAN specific pumps, or supplying pressure-fed duces expansion of the core gases and increases nozzle propellant. Engine system power balance analyses on pressure. 1,000,000-lbf-class LOX-hydrocarbon staged combus- Fig. 2 shows a cross-section of a nozzle downstream tion cycle engines with TAN have been performed. of the throat with the fuel and oxidizer injection ele- These analyzes indicate that providing the LOX and ments. Fig. 3 shows the axial pressure contour along the hydrocarbon fuels to the TAN injectors from the nozzle from throat to exit plane. This increased nozzle boost pump discharge is feasible with minimal im- pressure directly leads to increase thrust. The ignition pact on the main turbopump assembly (TPA), pre- source for the secondary propellant is the hot exhaust burners, and main chamber operating performance. from the core engine gases. Adding TAN-specific pumps would separate engine The TAN concept is scaleable to a wide range of and TAN development, and be the minimum impact of thrust-class engines from the very small thrust class of integrating a TAN subsystem into an existing engine 2000 to 1,500,000 lbf and larger. The TAN concept is system. applicable to various engine cycle schemes such as gas generator cycle, stage combustion cycles, and open and closed expander cycles. Primary Gas High Altitude Operation TAN OFF Primary Core Combustion TAN ON Fuel and Sea Level Operation Oxidizer Secondary Gas Reduces Injectors Secondary Expansion of Core and Combustion Increases Nozzle Pressure Fig. 1. TAN diagram for sea-level and high-altitude operation modes. Fig. 2. Schematic of TAN injection elements. S. Forde et al. / Acta Astronautica 59 (2006) 271–277 273 200 180 Augmentation ε = 25:1 160 Injector ε Location injection = 6:1 140 MRcore = 5.86 Pccore = 981 psia 120 MRTAN = 6.13 100 Augmentation On 80 60 40 Wall Static Pressure (psia) 20 Pambient No Augmentation 0 02413568 7910 Throat Axial Distance from Throat (inches) Fig. 3. Nozzle pressure as a function of axial distance from main thrust chamber throat with and without TAN. 3. TAN benefits 160 The TAN concept represents no less than a change in 150 the rocket propulsion paradigm and has wide ranging 140 benefits. This concept can create an engine with a higher thrust-to-weight ratio (T/W) for an engine, which can 130 be traded directly for increased payload. 120 AR=30 AR=40 NK-33 is a 350-klbf-class LOX-kerosene oxygen- AR=50 rich staged combustion cycle engine with one of the 110 AR=60 highest T/W of any booster engine. This engine was 100 used to show that an even higher T/W is achievable with Engine Sea Level Thrust/Weight 0 1020304060 TAN added to the engine. The sea-level T/W for the TAN% original NK-33 is 128:1. The T/W of a TAN equipped NK-33 can be increased to greater than 150:1 depend- Fig. 4. Sea-level thrust to weight vs. percent thrust augmentation on an NK-33 with TAN injector and TAN pumps. ing on the TAN augmentation, which is greater than a 17% improvement. Fig. 4 plots the NK-33 with TAN T/W as a function of the percent thrust augmentation with various nozzle holding the same nozzle exit pressure. Using a constant area ratios. Fig. 5 plots the weighted Isp (80% vacuum nozzle exit pressure can allow an NK-33 with a nozzle and 20% sea level). The 80/20 weighting was selected area ratio () of 58:1. A conventional NK-33 would have based on previous studies of a mission-averaged Isp separated flow with a 58:1 nozzle. for a two-stage-to-orbit launch vehicle. This 80/20 A similar comparison was performed using a LOX- weighting is dependent on many vehicle factors that hydrogen gas generator cycle engine in the 650-klbf- are mission dependent but is useful for demonstrating thrust class similar to an RS-68. The results have the potential vehicle benefits of a TAN-equipped engine same trends as the previous example; however, the where the TAN injection is throttled down or turned engine T/W improvement is even more dramatic. The off at an optimum altitude during the vehicle boost baseline LOX-hydrogen engine sea-level T/W was 46:1 phase. and can be improved to greater than 60:1, a greater than The two diagonal lines in Fig. 5 represent constant 60% improvement. Fig. 6 plots the LOX-H2 650-klbf pressure at the nozzle exit plane. The constant exit pres- gas generator cycle engine T/W vs. the percent thrust sure is one figure-of-merit for showing the benefits on an augmentation from TAN. Fig. 7 plots the weighted engine with TAN. The NK-33 with 40% thrust augmen- Isp (80% vacuum and 20% sea level) for the LOX tation yields a mission-averaged Isp gain of 4.5 s when hydrogen engine.
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