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 injecting 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

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. 274 S. Forde et al. / Acta Astronautica 59 (2006) 271–277

Weighted Isp vs %TAN 80%Isp(Vac)+20%∗Isp(SL) 333 Nozzle Attached Flow 332 Approx 6psi Exit Pressure Line 331 Nozzle Separated Flow 330 329 328

327 Constant Exit Pressure Line AR=30 As Baseline NK-33 Exit Pressure AR=40 326 Weighted Isp, Sec AR=50 AR=60 325 324 323 0102030 40 50 TAN %

Fig. 5. Weighted Isp vs. percentage of thrust augmentation of an NK-33 with TAN injector and TAN pumps.

70 in weight and complexity as compared to an engine gimbaled with hydraulic or electromagnetic actuators, thereby increasing propulsion reliability by providing 60 ﬁxed mounted engines. Possibly the most important beneﬁt of TAN is in- 50 creased engine system reliabilities resulting from oper- AR=21.5 ating the engine core at a chamber pressure below the

Engine T/W AR=30 AR=40 structural design condition and making up the required 40 AR=50 AR=60 thrust by operating TAN.

30 0 20 40 60 80 100 120 4. TAN status TAN% Two hot-fire test series have been conducted to date Fig. 6. Sea-level thrust to weight vs. percent thrust augmentation on 650-klbf LOX-hydrogen gas generator engine with TAN injector on the TAN concept to validate the design approach, and TAN pumps. combustion performance, and augmentation ratios. The first test series was performed in 2002 to validate the TAN concept with three tests having single propel- Other engine attributes that may benefit the launch lant injection into a 2000-lbf-thrust GOX/gaseous hy- vehicle include the higher thrust engine in the same drogen rocket engine. The main injector and chamber envelope as the conventional engine since the main was the same as used for the LANTR [5] testing with a turbopump and main injector is unchanged. This re- new TAN injection element added between the chamber duced engine packaging can reduce vehicle base area and the nozzle. Over 40% increase in sea-level thrust and drag, which is an important issue in fly-back boost- was demonstrated using a 25:1 expansion ratio nozzle ers. The TAN system can be used to compensate en- with the engine core operating at a chamber pressure of gine thrust in an engine-out condition for increased mis- 1000 psia. sion success and safe vehicle returns on multi-engine The second test series used the main GOX/gaseous vehicles. hydrogen injector from the LANTR program, and a Thrust vector control (TVC) is also possible by us- new water-cooled chamber, which incorporated a liquid ing TAN in an independently controlled circumferen- LOX-RP-1 TAN injector at an of 2:1 into a water- tially segmented configuration similar to the inert liq- cooled nozzle. This hardware has been successfully used uid injection that has been used on solid rocket motors. to demonstrate the liquid–liquid injection, mixing, and The TVC using TAN can result in significant reductions combustion of a heavy hydrocarbon fuel into the di- S. Forde et al. / Acta Astronautica 59 (2006) 271–277 275

Weighted Isp vs %TAN .8∗Isp(Vac)+.2∗Isp(SL) 416 Nozzle Attached Flow

414 Approximate 6 psia 412 Exit Pressure Line 410 408 Constant Exit Pressure Line for Typical GG LOX-Hydrogen Engine 406 Nozzle Separated Flow 404 AR=21.5

Weighted Isp, Sec 402 AR=30 400 AR=40 AR=50 398 AR=60 396 0 20 4060 80 100 120 TAN %

Fig. 7. Weighted Isp vs. percentage of thrust augmentation on 650-klbf LOX-hydrogen gas generator engine with TAN injector and TAN pumps.

Table 1 Test results of LOX-RP-1 TAN with GOX/H2 main injector

Test # Thrust Augmented Mass ﬂow Thrust (lbf) thrust (lbf) aug (%) aug (%)

105 1845 2112 24.9 14.5 106 1852 2126 25.8 14.8 107 1823 2331 41.3 27.8 108 1789 2512 55.1 40.4 109 1793 2376 45.7 32.5 110 1849 2331 37.2 26.1 111 1845 2345 36.4 27.1 112 1829 2546 57.0 39.2 113 1836 2405 43.2 31.0 116 1840 2514 51.7 36.7 117 1883 2577 53.7 36.9 Fig. 8. TAN in operation at the Sacramento A-Zone Test facility. 119 1854 2587 56.2 39.5 121 1836 2520 55.3 37.2 122 1292 1909 72.0 47.7 123 773 1371 108.1 77.4 combined cycle (RBCC) full-scale thruster [7] was per- 124 773 1217 77.1 57.4 formed where main chamber pressure, main chamber 126 759 1042 58.0 37.2 127 767 1114 63.3 45.3 mixture ratio, main chamber fuel film cooling, TAN to- 128 759 1037 56.5 36.6 tal flow rate, and TAN mixture ratio were varied. 129 767 1196 72.8 56.0 Table 1 briefly summarizes the tests. The testing 131 1822 2601 59.6 42.8 demonstrated C∗ efficiencies of 95% using conventional 132 1809 2537 56.1 40.2 truncated bell-shaped nozzle contours. Fig. 8 shows the 134 1805 2512 57.0 39.1 135 1809 2566 59.1 41.8 engine with TAN operating at Aerojet’s Sacramento 137 786 1314 87.9 67.1 A-Zone Test facility. The TAN injector after 28 tests, 138 783 1103 50.9 40.8 as shown in Fig. 9, was in excellent condition. Calculations of nozzle performance and wall pressures were made using a commercial computational fluid dynamics program. The results of calculations with vergent nozzle. The testing also demonstrated the dual and without thrust augmentation flow are compared with fuel capability of the TAN concept. Twenty-eight tests test data. There is good agreement between calculated using the LOX-RP-1 TAN on aerojet’s rocket based and measured quantities, including augmentation thrust. 276 S. Forde et al. / Acta Astronautica 59 (2006) 271–277

Fig. 11. Analysis results showing the Mach numbers with TAN operating.

thrusts, nozzle heat ﬂux characterization, and engine system and mission optimization trades. The next series of tests are planned at the 40-klbf-thrust class with LOX-hydrocarbon, either RP-1 or methane.

5. Conclusions

The TAN concept allows for obtaining high liftoff Fig. 9. TAN injector after 28 tests. thrust without increasing engine chamber pressure, resulting in a higher engine and vehicle T/W while increasing engine nozzle area expansion ratio capability for altitude performance beneﬁt. This results in an increase in mission Isp even with the higher thrust at takeoff. Other beneﬁts for TAN implementation are reducing the number of engines or de-rating the engine operating point, leading to increased reliability, low- ered cost, and improved vehicle packaging. Additional engine system test anchoring of analytical prediction models and design tools is underway. The next step is to work with the vehicle system designers to optimize the TAN system into current and future launch vehicle applications for increased vehicle capabilities.

References Fig. 10. Analysis results showing the Mach numbers without TAN operating. [1] M.J. Bulman, Aerojet-General Corporation, United States Patent US 6,568,171, Rocket vehicle thrust augmentation with divergent section of nozzle, May 27, 2003. [2] S.K. Borowski, R.R. Corban, D.W. Culver, M.J. Bulman, Fig. 10 shows the Mach numbers without TAN injec- M.C. McIlwain, A revolutionary lunar space transportation tion. Fig. 11 shows the results with TAN injection. The system architecture using extraterrestrial LOX-augmented NTR analytical models are currently being updated and val- Propulsion, AIAA-94-3343, 30th AIAA Joint Propulsion idated to the test data. Conference, July 1994. [3] M.J. Bulman, D.G. Messitt, T.M. Neill, S.K. Borowski, High area Further development of TAN is underway that will ratio LOX-augmented nuclear thermal rocket (LANTR) testing, address nozzle contour optimization, TAN injection AIAA-2000-3897, 36th AIAA Joint Propulsion Conference, starting and ending locations, TAN scaling to higher July 2000. S. Forde et al. / Acta Astronautica 59 (2006) 271–277 277

[4] M.J. Bulman, D.G. Messitt, T.M. Neill, S.K. Borowski, High area [6] M.J. Bulman, T.M. Neill, S.K. Borowski, LANTR engine system ratio LOX-augmented nuclear thermal rocket (LANTR) testing, integration, AIAA-2004-3864, 40th AIAA Joint Propulsion AIAA-2001-3369, 37th AIAA Joint Propulsion Conference, Conference, July 2004. July 2001. [7] T.M. Neill, J.D. Gladstone, Durability demonstration of small [5] M.J. Bulman, D.G. Messitt, T.M. Neill, S.K. Borowski, high performance rocket thrusters for RLV applications, AIAA- Continued LOX-augmented nuclear thermal rocket (LANTR) 99-2357, 35th AIAA Joint Propulsion Conference, June 1999. testing, AIAA-2002-3650, 38th AIAA Joint Propulsion Conference, July 2002.